Earth to orbit transportation system

ABSTRACT

Various embodiments of space launch vehicle systems and associated methods of manufacture and use are disclosed herein. In some embodiments, the systems include a reusable, horizontal takeoff/horizontal landing (HTHL), ground-assisted single-stage-to-orbit (SSTO) spaceplane that is capable of providing frequent deliveries of people and/or cargo to Low Earth Orbit (LEO). In some embodiments, the spaceplane can takeoff with the aid of a rocket-powered sled that, in addition to providing additional thrust for takeoff, can also provide propellant for the spaceplane engines during the takeoff run so that the spaceplane launches with full propellant tanks.

CROSS-REFERENCE TO RELATED APPLICATIONS INCORPORATED BY REFERENCE

The present application is a continuation of International PatentApplication No. PCT/US19/34003, filed May 24, 2019, and titled EARTH TOORBIT TRANSPORTATION SYSTEM, which claims priority to U.S. ProvisionalPatent Application No. 62/676,809, filed May 25, 2018, and titled EARTHTO ORBIT TRANSPORTATION SYSTEM, each of which is incorporated herein byreference in its entirety.

TECHNICAL FIELD

The present disclosure is generally related to vehicles and associatedsystems and methods for transporting crew and cargo to space (e.g.,reusable Earth-to-orbit vehicles).

BACKGROUND

There are a number of existing launch vehicles available fortransporting crew and cargo to Low Earth Orbit (LEO) and to existingin-orbit systems, such as the International Space Station (ISS). TheSpace Shuttle and the SpaceX Falcon 9 are two such vehicles. The SpaceShuttle, however, was costly to operate, and although many of itssystems were reusable, with the notable exception of the large externaltank, they required a great deal of logistics support to refurbish,reassemble, and relaunch. Additionally, both the Space Shuttle and theFalcon 9 were designed to carry relatively heavy payloads of about50,000 lbs. to LEO. As a result, these vehicles do not present viable,relatively low-cost options for transporting crew and/or lighter cargo(e.g., about 5-10,000 lbs.) to LEO.

Space transportation systems include single-stage-to-orbit (SSTO) launchvehicles as well as multi-stage-to-orbit vehicles. In the early 1970's,Boeing developed a design proposal for a REUSABLE AERODYNAMIC SPACEVEHICLE (RASV). Although the RASV was never built, the proposed designwas a SSTO Horizontal Takeoff, Horizontal Landing (HTHL) spaceplane thatutilized a sled boost assisted launch. The vehicle was primarilydirected toward military space missions, and utilized a welded metalhoneycomb airframe with integral thermal protection. The proposedpropulsion system included very complex, high maintenance liquid oxygen(LOX)/liquid hydrogen (LH2) Space Shuttle main engines.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1A is a partially schematic, top rear isometric view of anaerospace vehicle configured in accordance with embodiments of thepresent technology, and FIGS. 1B-1E are series of front, side, top frontisometric, and bottom front isometric views, respectively, of theaerospace vehicle.

FIG. 2 is a partially schematic, partially cross-sectioned top view ofthe aerospace vehicle configured in accordance with embodiments of thepresent technology.

FIG. 3A is a cross-sectional side isometric view illustrating a forwardportion of the fuselage of the aerospace vehicle, and FIG. 3B is anisometric view of the aerospace vehicle docking with an orbiting stationin space, in accordance with embodiments of the present technology.

FIGS. 4A-4C are a series of top rear isometric, top cross-sectionalisometric, and side cross-sectional isometric views, respectively, of arocket powered launch sled configured in accordance with embodiments ofthe present technology, and FIG. 4D is a schematic diagram illustratingpropellant distribution between the launch sled and the vehicle inaccordance with an embodiment of the present technology.

FIGS. 5A and 5B are partially schematic side and end views,respectively, of a coupling for movably mounting a launch sled to one ormore launch rails in accordance with embodiments of the presenttechnology.

FIG. 6 is a flow diagram of a method for attaching the aerospace vehicleof FIGS. 1A-1E to the launch sled of FIGS. 4A-4C in accordance withembodiments of the present technology.

FIG. 7A is a front isometric view of the aerospace vehicle operablymounted to the launch sled, and FIGS. 7B-7E are a series of side viewsillustrating various stages of a process for attaching the vehicle tothe launch sled and preparing the vehicle for launch in accordance withembodiments of the present technology.

FIG. 8A is a partially schematic side view illustrating forward supportarm and electrical umbilical connections to an aerospace vehicle, andFIG. 8B is a partially schematic rear view illustrating rear support armand propellant umbilical connections to the aerospace vehicle,configured in accordance with embodiments of the present technology.

FIGS. 9A and 9B are side and rear views, respectively, of a support armhold and release mechanism in a first stage of operation, and FIG. 9C isa side view of the support arm hold and release mechanism in a secondstage of operation, in accordance with embodiments of the presenttechnology.

FIGS. 10A and 10B are a front view and a side cross-sectional view,respectively, of a vehicle support arm interface configured inaccordance with embodiments of the present technology.

FIGS. 11A and 11B are a front view and a side cross-sectional view,respectively, of a support arm end fitting configured in accordance withembodiments of the present technology.

FIGS. 12A and 12B are side and rear views, respectively, of anothersupport arm hold and release mechanism in a first stage of operation,and FIGS. 12C and 12D are side and rear views, respectively, of thesupport arm hold and release mechanism in a second stage of operation,in accordance with embodiments of the present technology.

FIG. 13 is a schematic diagram of an aerospace vehicle and a launch sledoperably coupled to a propellant management system configured inaccordance with embodiments of the present technology.

FIG. 14 is a flow diagram of a routine for loading propellants onto alaunch sled and an aerospace vehicle in accordance with embodiments ofthe present technology.

FIG. 15 is a schematic diagram of a system architecture for controllinga propellant management system, a launch sled, and an aerospace vehicleduring propellant loading, in accordance with embodiments of the presenttechnology.

FIG. 16A is a flow diagram of a routine for operating an aerospacevehicle and a launch sled during a takeoff run; FIG. 16B is a flowdiagram of a routine for confirming safe liftoff conditions prior torelease of the aerospace vehicle from the launch sled; FIG. 16C is aflow diagram of a routine for liftoff of the aerospace vehicle from thelaunch sled; and FIG. 16D is a flow diagram of a routine for abortingliftoff of the aerospace vehicle from the launch sled, in accordancewith embodiments of the present technology.

FIGS. 17A-17D are a series of schematic diagrams illustrating operationof an aerospace vehicle and a launch sled at various stages of a launchprocess in accordance with embodiments of the present technology.

FIG. 18A is a schematic diagram of a system architecture for controllingan aerospace vehicle and a launch sled after separation from apropellant management system; and FIG. 18B is a schematic diagram of asystem architecture for controlling the aerospace vehicle after liftofffrom the launch sled, in accordance with embodiments of the presenttechnology.

FIG. 19 is a partially schematic diagram illustrating various stages ofa flight sequence of an aerospace vehicle in accordance with embodimentsof the present technology.

FIG. 20 is a flow diagram of a routine for performing an ascent of anaerospace vehicle in accordance with embodiments of the presenttechnology.

FIG. 21 is a chart listing example types of mission aborts and enginefailures/degradations that an aerospace vehicle could experience, inaccordance with embodiments of the present technology.

FIG. 22 is a flow diagram of a routine for responding to an engineanomaly after liftoff of an aerospace vehicle in accordance withembodiments of the present technology.

FIG. 23A is a schematic diagram of a control system for an aerospacevehicle, and FIG. 23B is a schematic diagram of a control system for alaunch sled, configured in accordance with embodiments of the presenttechnology.

FIG. 24A is a partially schematic diagram of an oxidizer tank configuredin accordance with embodiments of the present technology, and FIG. 24Bpresents a graph of various pressures versus time for the oxidizer tankand its environment.

FIG. 25A is a partially schematic diagram of a fuel tank configured inaccordance with embodiments of the present technology, and FIG. 25Bpresents a graph of various pressures versus time for the fuel tank andits environment.

FIGS. 26A and 26B are top front isometric and bottom front isometricviews, respectively, of the aerospace vehicle of FIGS. 1A-1Dillustrating various aspects of the airframe and an associated thermalprotection system, configured in accordance with embodiments of thepresent technology.

DETAILED DESCRIPTION

The following disclosure describes various embodiments of space launchvehicle systems and associated methods of manufacture and use. In someembodiments, the systems include a fully reusable, horizontaltakeoff/horizontal landing (HTHL), ground-assisted single-stage-to-orbit(SSTO) spaceplane that is capable of providing frequent deliveries ofpeople and/or cargo to Low Earth Orbit (LEO). As described in greaterdetail below, the spaceplane can take off with the aid of arocket-powered sled that, in addition to providing thrust for takeoff,can also provide propellant for the spaceplane engines during thetakeoff run so that the spaceplane launches with full propellant tanks.In some embodiments, the sled can utilize magnetic levitation andmagnetic propulsion to provide thrust for takeoff. After several hoursor days in orbit, the spaceplane can fly back to Earth and land on aconventional runway having a length of, for example, about 10,000 ft.Embodiments of the systems disclosed herein can enable the expansion ofthe existing space industry by providing low-cost access to orbitaldestinations, such as the International Space Station (ISS), for peopleand light to medium cargo (e.g., about 5,000 lbs.) at relatively highfrequency (e.g., as often as twice a week).

Certain details are set forth in the following description and in FIGS.1-26B to provide a thorough understanding of various embodiments of thepresent technology. In other instances, well-known structures,materials, operations and/or systems often associated with aerospacevehicle structures, propulsion systems, control systems, flightsequences, control routines, etc. are not shown or described in detailin the following disclosure to avoid unnecessarily obscuring thedescription of the various embodiments of the technology. Those ofordinary skill in the art will recognize, however, that the presenttechnology can be practiced without one or more of the details set forthherein, or with other structures, methods, components, and so forth.

The accompanying Figures depict embodiments of the present technologyand, unless otherwise specified, are not intended to be limiting of itsscope. The sizes of various depicted elements are not necessarily drawnto scale, and these various elements may be arbitrarily enlarged toimprove legibility. Component details may be abstracted in the Figuresto exclude details such as position of components and certain preciseconnections between such components when such details are unnecessaryfor a complete understanding of how to make and use the disclosedtechnology. Many of the details, dimensions, angles and other featuresshown in the Figures are merely illustrative of particular embodimentsof the disclosure. Accordingly, other embodiments can have otherdetails, dimensions, angles and features without departing from thepresent technology. In addition, those of ordinary skill in the art willappreciate that further embodiments of the present technology can bepracticed without several of the details described below.

In general, identical reference numbers in the Figures identifyidentical, or at least generally similar, elements. To facilitate thediscussion of any particular element, the most significant digit ordigits of any reference number refers to the Figure in which thatelement is first introduced. For example, element 110 is firstintroduced and discussed with reference to FIG. 1.

FIG. 1A is a partially schematic, top rear isometric view of anaerospace vehicle 100 (which can also be referred to as a spaceplane)configured in accordance with embodiments of the present technology.FIGS. 1B-1E are front, side, top front isometric, and bottom frontisometric views, respectively, of the vehicle 100. Referring first toFIGS. 1A-1C, in the illustrated embodiment the vehicle 100 is anHTHL/SSTO vehicle having a pair of highly swept wings 104 (identifiedindividually as a left wing 104 a and a right wing 104 b) extendingoutwardly from a fuselage 102 to provide lift during flight in theEarth's atmosphere. The trailing edge portion of each of the wings 104a, b includes a corresponding elevon 106 a, b for vehicle pitch and rollcontrol. Additionally, the vehicle 100 includes a pair of verticalstabilizers 110 (identified individually as a left vertical stabilizer110 a and a right vertical stabilizer 110 b) having correspondingrudders 108 a, b on trailing edge portions thereof for providing thevehicle 100 with yaw control. The fuselage 102 can include a door 103 inan upper portion of a crew cabin 112 for crew ingress and egress.Additionally, in the illustrated embodiment the forward portion of thefuselage 102 includes a movable hatch 114 for providing access to adocking port (not shown in FIG. 1A-1E) for docking the vehicle 100 withan on-orbit station, such as the ISS, and enabling human and/or cargomovement therebetween.

As shown in FIG. 1A, in the illustrated embodiment the aft portion ofthe fuselage 102 carries a propulsion system 111 having one or morerocket engines 120 (three are shown in FIG. 1A; identified individuallyas first, second, and third rocket engines 120 a-c, respectively). Eachof the engines 120 a-c has a corresponding nozzle 117 positionedgenerally proximate the trailing edge portion of the wings 104 a, bbetween the vertical stabilizers 110 a, b. As described in greaterdetail below, in some embodiments the rocket engines 120 a-c areconfigured to burn liquid oxygen (LOX) and jet fuel as propellants. Thejet fuel can include common kerosene-types of aviation fuel designed foruse in aircraft powered by gas-turbine engines including, for example,“Jet-A.” Additionally, in some embodiments the engines 120 a-c caninclude dual area ratio nozzles with injection ports for tripping theexhaust flow. When the injection ports are inactive, the exhaust flowoccupies the entire cross-sectional area at the exit plane of the nozzle117, producing a first effective nozzle area. The area ratio of thefirst effective nozzle area to the area at the nozzle throat can berelatively large, which is suitable for high altitude performance. Forexample, the area ratio can be about 60:1 in some embodiments. When theinjection ports are activated, the flow from the injection ports tripsthe exhaust flow and produces a shockwave that limits the effective flowover of the nozzle 117 to a smaller, second effective nozzle area. Thearea ratio of the second effective nozzle area to the area at the nozzlethroat can be relatively small, for example, about 33:1. Accordingly,since the nozzle flow is typically over-expanded at low altitude (as acompromise to improve high altitude performance), the nozzle exit areareduction provided by the tripped exhaust flow can improve nozzleefficiency at low altitude. In some embodiments, the engines 120 a-c canbe at least generally similar in structure and function to enginesdescribed in U.S. Provisional Application No. 62/693,829, filed on Jul.3, 2018, and titled “ROCKET PROPULSION SYSTEMS AND ASSOCIATED METHODS,”which is incorporated herein by reference in its entirety.

The vehicle 100 can further include orbital maneuvering system (OMS)engines 122 (identified individually as a first OMS engine 122 a and asecond OMS engine 122 b) having nozzles positioned just above thenozzles for the main engines 120 a-c. In some embodiments, the OMSengines 122 can be bipropellant rocket engines that use LOX andcompressed natural gas (CNG; consisting mostly of methane). The use ofLOX and CNG provides a gas-gas propellant solution that can be used in ablowdown system that relies on gas pressure to drive the propellantsinto the OMS engines 122. As described in greater detail below, the OMSengines provide steering and directional control when the vehicle 100 isin space, and can enable the vehicle 100 to reorient in space fordeorbiting and reentry into the Earth's atmosphere. Although theillustrated embodiment of the vehicle 100 includes three main engines120 that use LOX and Jet-A as propellants, the technology disclosedherein is not limited to any particular number of engines or anyparticular types of propellants. Accordingly, it will be understood thatvehicles configured in accordance with the present technology caninclude more or fewer engines using other types of propellants (e.g.,LOX and refined petroleum (e.g., RP-1), LOX/liquid hydrogen, LOX/CNG,etc.) consistent with the present disclosure.

The vehicle 100 can include a controller 140 having one or moreprocessors 142 that can control various operations and functions of thevehicle 100 in accordance with computer-readable instructions stored onsystem memory 144. The controller 140 can receive inputs 113 and issueoutputs 115. By way of example, the inputs 113 can include controlsignals and commands from, e.g., ground systems, the crew, etc.; flightparameters such as airspeed and/or ground speed, altitude, dynamicpressure, temperature, etc.; engine operating parameters; propellantparameters; vehicle positional and directional information; etc. Theoutputs 115 can include commands directing vehicle operation, includingcontrol surface operation via associated valves, actuators, and/or othercomponents; engine operation including start, stop, and throttlesettings; data and telemetry transmissions; etc. The processor 142 caninclude any logic processing unit, such as one or more centralprocessing units (CPUs), digital signal processors (DSPs), acceleratedprocessing units (APUs), application-specific integrated circuits(ASICs), etc. The processor 142 may be a single processing unit ormultiple processing units distributed across multiple systems and/orsubsystems of the vehicle 100. The processor 142 is operably connectedto the memory 144 and may be operably connected to various systems ofthe vehicle 100 to transmit instructions and/or receive input therefrom.The memory 144 can include read-only memory (ROM) and random-accessmemory (RAM) or other storage devices that store executableapplications, test software, databases and/or other software requiredto, for example, control or at least partially control the flight,propulsion, power, avionics, telemetry, environmental, and/or othersystems of the vehicle 100 in accordance with the methods describedherein, and enable the vehicle 100, its systems and occupants tocommunicate and/or exchange data and information with remote computers(e.g., computers on Earth and/or in orbit) and/or other devices.

In some embodiments, the vehicle 100 includes all of the systemsnecessary for implementing the mission sequences described herein. Suchsystems can include, for example, a communications system 146 for, e.g.,wireless communications (including crew communications, digitalcommunications between processing devices, etc.) between the vehicle 100and, e.g., ground control, ground stations, orbiting stations, etc. Thecommunication system 146 can include, for example, wirelesstransceivers, antennae, etc. for broadcasting transmissions to andreceiving transmissions from remote locations. The vehicle systems canalso include an electrical power and distribution system 148; anavigation system 150; a flight controls system 152 for affectingactuation of the vehicle control surfaces, engine throttles, landinggear, etc.; avionics 154; a hydraulic system 156 for, e.g., controlsurface and landing gear actuation; and an environmental control system158 for maintaining, e.g., air conditioning, etc. for human occupancy.The foregoing systems are non-exclusive, and it will be understood thatsome embodiments of the vehicle 100 can include other control andoperating systems, while other embodiments of the vehicle 100 may notinclude one or more of these systems.

As shown in FIGS. 1B and 1C, in the illustrated embodiment the vehicle100 includes a landing gear system having a nose gear 126 a, a left maingear 126 b, and a right main gear 126 c. (In FIG. 1E, the landing gear126 a-c are retracted and stowed behind corresponding gear doors 129a-c.) As described in more detail below, the vehicle 100 takes off withthe assistance of a launch sled and, as a result, the landing gear 126a-c are retracted and stowed into associated gear bays during takeoff.For landing, the landing gear 126 a-c are deployed in a manner that isat least generally similar to conventional commercial aircraft. Sincethe landing gear 126 a-c are only designed to carry the loads associatedwith landing the vehicle 100 when it is not carrying a full load ofpropellant and is therefore relatively light, the landing gear 126 a-ccan be substantially lighter than they would otherwise be if they weredesigned to support the vehicle 100 during takeoff with a full load ofpropellant. This weight savings results in an increased payloadcapacity.

Referring next to FIG. 1D, in some embodiments the vehicle 100 caninclude an antenna 134 which can be deployed from the fuselage 102 oncethe vehicle 100 is in orbit to facilitate communications between Earth,on-orbit stations, and/or other remote locations. The vehicle 100 canfurther include a plurality of thrusters 136 positioned at variouslocations on the exterior of the vehicle 100 to provide attitude controlwhile on orbit. Such reaction control system (RCS) thrusters caninclude, for example, relatively small monopropellant thrusters known inthe art. In some embodiments, the vehicle 100 may be configured to useonly “green” propellants. In such embodiments, the thrusters 136 can usea hydroxylammonium nitrate-based propellant known as AF-M315E, and/or apropellant known as LMP-103S, which is based on the oxidizer ammoniumdinitramide. Both of these propellants are less toxic than, for example,hydrazine. In other embodiments, the thrusters 136 can use hydrazine ina conventional manner. In some embodiments, the vehicle 100 can includea pair of thrusters 136 a, b toward an upper aft portion of the fuselage102, and a pair of thrusters 136 c, d near the tip portion of each ofthe wings 104 a, b. Additionally, the vehicle 100 can also include apair of thrusters 136 e, f at the base of each vertical stabilizer 110a, b, a pair of thrusters 136 g, h toward an upper forward portion ofthe fuselage 102 just behind the crew cabin 112, and another pair ofthrusters 136 i, j on opposite sides of the fuselage near the samelocation. As shown in FIG. 1E, additional thrusters 136 k, l and 136 m,n can also be included on the underside of the vehicle 100 proximate theforward and aft portions the fuselage 102, respectively. Selectiveactivation of the various thrusters 136 enable the vehicle to bepositioned in virtually any attitude while in orbit to facilitate, forexample, docking with on-orbit stations, transmission and/or receptionof communications, planetary viewing, etc. The thruster locationsillustrated in FIGS. 1D and 1E are provided by way of examples of someembodiments. Accordingly, it will be understood that other embodimentscan have more or fewer thrusters and/or thrusters in other locations.

As also shown in FIG. 1E, in the illustrated embodiment the vehicle 100further includes propellant interfaces 124 (identified individually as afirst propellant inlet interface 124 a and a second propellant inletinterface 124 b), and support arm interfaces 127 (identifiedindividually as a forward support arm interface 127 a, a first aftsupport arm interface 127 b, and a second aft support arm interface 127c). Each of the support arm interfaces 127 a-c can include a couplingconfigured to releasably engage a corresponding support arm for mountingthe vehicle 100 to the launch sled (not shown in FIG. 1E) prior to andduring takeoff. In some embodiments, each of the propellant inletinterfaces 124 a, b is positioned just inboard of (and laterallyadjacent to) the corresponding support arm interface 127 b, c. Asdescribed in greater detail below, each of the propellant interfaces 124a, b can include a quick-disconnect valve and/or other suitable couplingfor releasably connecting a corresponding propellant umbilical (e.g., apropellant conduit; which can also be referred to as a propellant line)from the launch sled to the vehicle 100 and sealing the interface whenthe propellant umbilical is disconnected. Additionally, the propellantinlet interfaces 124 a, b can also include doors that close flush withthe outer surface of the fuselage 102 to protect the interfaces 124 a, bfrom aerothermal heating, etc. Propellants (e.g., LOX and Jet-A) aretransferred from the launch sled to the vehicle 100 via the propellantumbilicals and the inlet interfaces 124 a, b for operation of the mainvehicle engines 120 a-c during takeoff. For example, in some embodimentsthe first propellant inlet interface 124 a can be configured to receiveLOX from the launch sled via a LOX umbilical, and the second propellantinlet interface 124 b can be configured to receive fuel from the launchsled via a fuel umbilical. Additionally, as described in greater detailbelow, in some embodiments the first propellant inlet interface 124 acan also be configured to recirculate vented/warmed LOX from the vehicle100 back to the launch sled via the LOX umbilical, and the secondpropellant inlet interface 124 b can also be configured to recirculatevented fuel from the vehicle 100 back to the launch sled via the fuelumbilical.

In addition to the propellant inlet interfaces 124 a, b, the vehicle 100further includes an electrical interface 125 positioned just aft of theforward support arm interface 127 a. The electrical interface 125 isconfigured to releasably connect to an electrical umbilical that extendsfrom the launch sled to the vehicle 100. As described in greater detailbelow, the electrical interface 125 can include one or more electricalreceptacles configured to receive one or more corresponding connectorson the electrical umbilical to enable transmission of commands, power,and data between the vehicle 100 and the launch sled. Like thepropellant interfaces 124 a, b, the electrical interface 125 can alsoinclude a door that closes flush with the outer surface of the fuselage102 after the electrical umbilical has been disconnected to protect theinterface 125 from aerothermal heating, etc.

FIG. 2 is a partially schematic, partially cross-sectioned top view ofthe vehicle 100 configured in accordance with embodiments of the presenttechnology. In one aspect of this embodiment, the oxidizer (e.g., LOX)for the main engines 120 a-c is contained in a fuselage tank 242 that isformed by external sidewalls of the fuselage 102 and, accordingly, thetank 242 can have a cross-sectional shape that follows thecross-sectional shape (contour) of the fuselage 102. More specifically,as discussed in greater detail below, in some embodiments the LOX issubcooled to a temperature of, for example, about −320 degreesFahrenheit (F) (i.e., about −196 degrees Celsius (C)). At thistemperature, the LOX vapor pressure is sufficiently low that thepressure differential across the walls of the fuselage tank 242 is lessthan about 3 psig, such as between 2-3 psig. By maintaining the tankpressure at about 2-3 psig, the structural loads on the tank walls arerelatively low. As a result, the tank 242 does not have to have theshape of a conventional high-pressure propellant tank or pressure vessel(e.g., a spherical shape or a cylindrical shape having a circularcross-section). This enables the internal volumes of the airframe (e.g.,the fuselage 102) to be used as a LOX tank, while at the same time beingshaped for optimum aerodynamic performance without requiring anystructural reinforcement to accommodate high tank pressure loads. Forexample, in some embodiments the oxidizer tank 242 can have across-sectional shape that is non-circular, such as an oval, or nearoval, cross-sectional shape, an elliptical cross-sectional shape, anasymmetric cross-sectional shape, and/or other non-circularcross-sectional shapes. Additionally, as shown by the plan view of FIG.2, in some embodiments the cross-section of the oxidizer tank 242 canvary in both shape and/or size along the length of the fuselage 102. Inother embodiments, however, the oxidizer tank 242 can have other shapes,such as cylindrical and/or spherical shapes. The oxidizer tank 242 isconnected in fluid communication with the first propellant inletinterface 124 a (FIG. 1E). In another aspect of this embodiment, each ofthe wings 104 a, b includes a corresponding fuel tank 240 a, b thatcontains the fuel for the vehicle main engines 120 a-c. The fuel tanks240 a, b fill much of the interior volumes of the wings 104 a, b in theforward strake regions and the main wing regions, except for the volumesproximate wing leading edge regions 246 and main spar sections 248. Thefuel tanks 240 a, b are connected in fluid communication with the secondpropellant inlet interface 124 b (FIG. 1E). The OMS propellant tanks 244(e.g., LOX and CNG tanks) are positioned between the main engineoxidizer tank 242 and the main engines 120 a-c.

FIG. 3A is a side cross-sectional view of the forward portion of thefuselage 102 illustrating the crew cabin 112 and an adjacent payload bay374. In some embodiments, the payload bay 374 can include an airlock anda docking port 358. In other embodiments, in addition to or instead ofthe docking port 358, the payload bay 374 can include a payload supportand deployment system (not shown) configured to carry payloads (e.g.,one or more satellites) and deploy them into orbit. FIG. 3B is anisometric view of an orbiting station 368 operably coupled to thevehicle 100 via the docking port 358 in accordance with embodiments ofthe present technology. Referring first to FIG. 3A, in the illustratedembodiment the crew cabin 112 can include a plurality of seats 350(e.g., 5 seats) for vehicle crew and/or passengers. The crew cabin 112can also include a plurality of windows 370 for occupant viewing outsideof the vehicle 100.

In some embodiments, the crew cabin 112 is a self-contained unit thatcan separate from the rest of the fuselage 102 in the event of a missioncritical failure that occurs at any point during flight. Morespecifically, an aft portion of that crew cabin 112 can be sealablyenclosed by a pressure bulkhead 354 that enables the crew cabin 112 tomaintain internal pressure during all phases of operation. Additionally,the crew cabin 112 can be structurally attached to the rest of thefuselage 102 by a frangible joint 356 that extends around thecircumference of the fuselage just aft of the bulkhead 354. Thefrangible joint 356 can include a pyrotechnically actuated explosivedevice (e.g., Super Zip from Ensign-Bickford Aerospace Company) orlinear shaped charge that structurally attaches the crew cabin 112 tothe fuselage 102 until actuated in response to a separation signal. Uponactuation, the frangible joint 356 breaks to immediately detach the crewcabin 112 from the rest of the fuselage 102. In other embodiments,instead of (or in addition to) the frangible joint 356, the crew cabin112 can be attached to the rest of the fuselage 102 with a plurality ofexplosive bolts and/or other known separating devices to enable the crewcabin 112 to be quickly disengaged and separated from the fuselage 102in the event of a mission critical failure of one or more vehiclesystems.

The vehicle 100 can include a number of subsystems to facilitateseparation of the crew cabin 112 from the rest of the fuselage 102 andsafe return of the crew cabin 112 to Earth in the event of a missioncritical failure. For example, in the illustrated embodiment the crewcabin 112 can include a recovery chute 364 positioned toward an upperaft portion of the crew cabin 112, and a downward-firing thruster 360 aand a first aft-firing thruster 360 b positioned toward a lower aftportion of the crew cabin 112. Although not shown by virtue of thesection view, a second aft-firing thruster is positioned adjacent to thefirst aft-firing thruster 360 b on the opposite sided of the vehiclecenterline. The thrusters 360 a, b can be conventional bipropellant orgreen propellant thrusters that receive propellant from correspondingfuel and oxidizer tanks 362 positioned beneath a floor 372 of the crewcabin 112. The nozzle of the first thruster 360 a can be directedgenerally downward and aft, and the nozzle of the second thruster 360 bcan be directed generally aft. The recovery chute 364 can include one ormore parachutes that are deployed from the crew cabin 112 afterseparation from the rest of the fuselage 102 and when the crew cabin 112is at an appropriate altitude on descent.

As described in greater detail below, in the event the vehicle 100experiences a mission critical failure at any point during flight, thefrangible joint 356 can be activated to separate the crew cabin 112 fromthe rest of the fuselage 102. Immediately after separation, thethrusters 360 a, b can be ignited to quickly move the crew cabin 112 asafe distance away from the rest of the vehicle 100. Once the crew cabin112 descends to an appropriate altitude, the recovery chute 364 can bedeployed to slow the decent through the Earth's atmosphere. The recoverychute 364 can be positioned to properly orient the crew cabin 112 duringdescent through the atmosphere. Additionally, the crew cabin 112 canalso include a thermal protection system (e.g., a fibrous, reinforcedoxidation-resistant composite covering) on at least the forward-facingsurfaces to provide sufficient heat shielding during reentry.

In another aspect of this embodiment, the crew cabin includes a passage352 (with a sealable door, not shown) that enables the crew of thevehicle 100 to move back and forth between the crew cabin 112 and theadjacent airlock or payload bay 374 when in orbit. As shown in FIG. 3B,in operation the hatch 114 can be opened to enable the docking port 358to sealably engage and structurally attach to a corresponding dockingport 366 on the orbiting station 368. Once properly docked, crew and/orcargo from the vehicle 100 can move back and forth between the vehicle100 and the orbiting station 368.

FIG. 4A is a partially schematic isometric view of a launch sled 400configured in accordance with embodiments of the present technology, andFIGS. 4B and 4C are top cross-sectional isometric and sidecross-sectional isometric views, respectively, of the launch sled 400.Referring first to FIG. 4A, the sled 400 runs on three heavy duty rails410 a-c that react the combined loads from the sled 400 and the vehicle100 during acceleration and deceleration. In some embodiments, the rails410 a-c can be about two miles long. In the illustrated embodiment, thesled 400 includes a body or chassis 402 having a center section 446 a, aleft outer section 446 b, and a right outer section 446 c that areconnected by an aerodynamic top plate 406. The sections 446 a-c can alsobe referred to as “trucks.” The underside of each of the sections 446a-c is moveably engaged with a corresponding rail 410 a-c by at least aforward coupling 454 a and an aft coupling 454 b. In operation, thecouplings 454 a, b enable the sled to move fore and aft on the rails 410a-c, while keeping the sled 400 attached to the rails.

The sled 400 further includes a forward support arm 416 a operablycoupled to a forward portion of the center section 446 a, and first andsecond aft support arms 416 b and 416 c, respectively, operably coupledto the first outer section 446 b and the second outer section 446 c,respectively. In some embodiments, the proximal end portion of each ofthe support arms 416 a-c is pivotably coupled to the sled chassis 402,and the arms 416 a-c are operably coupled to a drive system (e.g., anelectromechanical system, a hydraulic system, a pneumatic system, etc.;not shown) that moves (e.g., rotates) the arms 416 a-c through theirranges of operating motion. The distal end portion of each support arm416 a-c is configured to be releasably attached to the correspondingsupport arm interface 127 a-c on the vehicle 100 by means of a hold andrelease mechanism. For example, in some embodiments the distal endportion of each support arm 416 a-c can include a fitting 422 (e.g., aball fitting or other suitable fitting that can permit rotation of thearm 416 while the vehicle 100 is attached) that is configured to bereleasably engaged with the corresponding support arm interface 127 a-c(e.g., a ball socket) by the hold and release mechanism. In someembodiments, the hold and release mechanism can include a mechanicalclamp mechanism that holds the distal end portion of each control arm416 a-c to the corresponding interface 127 a-c until commanded torelease. In other embodiments, other suitable hold and releasemechanisms known in the art can be used to releasably attach the supportarms 416 a-c to the corresponding interfaces 127 a-c.

In FIG. 4A, each of the support arms 416 a-c is illustrated in threedifferent operational positions: a lowermost or stowed position, anintermediate position in which the arm is rotated partially upward toengage and raise the vehicle 100 for stowage of the landing gear 126a-c, and a launch position in with the arm is rotated fully aft and/orupward to optimize (or nearly optimize) vehicle angle of attack forseparation and lift off. The support arms 416 a-c react all theacceleration and deceleration loads between the vehicle 100 and the sled400 during takeoff, thereby eliminating the need for the vehicle 100 toinclude a landing gear system that is rated for takeoff loads.Additionally, the positioning of the support arms 416 a-c can beadjustable to optimize the vehicle angle of attack as needed for eachmission. Although the illustrated embodiment includes three support arms416 a-c, in other embodiments the sled 400 can include more support arms(e.g., four, five, or more arms) or fewer arms (e.g., two or one arm).

In the illustrated embodiment, the sled 400 includes three rocketengines 404 (identified individually as rocket engines 404 a-c) whichare mounted to the aft portions of the corresponding sled sections 446a-c. In other embodiments, the sled 400 can have more or fewer rocketengines (e.g., one to five or more rocket engines). The rocket engines404 a-c have nozzles with area ratios optimized for ground performance.Like the main vehicle engines 120 a-c (FIG. 1A), the engines 404 a-c canbe bipropellant rocket engines configured to burn, for example, LOX andJet-A. Since the engines 404 a-c operate at a constant or near-constantaltitude, their nozzles can be configured for single area ratiooperation and optimum performance at the altitude at which the sledoperates (e.g., sea level or near sea level). In some embodiments, theengines 404 a-c can be at least generally similar in structure andfunction to engines described in U.S. Provisional Application No.62/693,829, filed on Jul. 3, 2018, and titled “ROCKET PROPULSION SYSTEMSAND ASSOCIATED METHODS,” which is incorporated herein by reference inits entirety.

In some embodiments, each of the engines 404 a-c can have dedicatedpropellant tanks (e.g., a dedicated oxidizer (e.g., LOX) tank 452 and/ora dedicated fuel (e.g., Jet-A) tank 450) mounted and housed within theenclosure of the corresponding sled section 446 a-c, as shown in FIGS.4B and 4C. The propellant tanks 450 and 452 are sized and configured toprovide propellant to the corresponding engines 404 a-c for the durationof the sled run. In other embodiments, the sled 400 can carry a singleoxidizer tank and/or a single fuel tank that provides sufficientpropellant for all the engines 404 a-c. In further embodiments, all or aportion of the engines 404 a-c may use different propellants, and/or oneor more of the engines 404 a-c may be solid rocket motors.

In addition to the sled propellant tanks 450 and 452, the sled 400 canalso carry auxiliary propellant tanks (e.g., an oxidizer (e.g., LOX)tank 458 and a fuel (e.g., Jet-A) tank 456; shown in cross-section inFIGS. 4B and 4C) for providing propellant to the vehicle engines 120 a-c(FIG. 1A) for the duration of the sled run. As described in greaterdetail below, this assures that the vehicle 100 is as full of propellantas possible at liftoff from the sled, thereby minimizing (or at leastgreatly reducing) the vehicle dry mass penalty at lift off. In someembodiments, the auxiliary propellant tanks 456 and 458 can be housed inthe center section 446 a. In other embodiments, the propellant tanks 456and 458 can be carried in one or both of the left outer section 446 band/or the right outer section 446 c. In other embodiments, one or bothof the auxiliary propellant tanks 456 and 458 can be omitted, and one ormore of the sled propellant tanks 452 and 450 can be sized andconfigured to provide propellant to both the corresponding sledengine(s) 404 as well as the vehicle engines 120 a-c for the duration ofthe sled run.

The sled 400 further includes a first propellant outlet interface 414 aand a second propellant outlet interface 414 b located on or proximatethe top plate 406. In the illustrated embodiment, the first propellantoutlet interface 414 a is positioned inboard of, and laterally adjacentto, the base of the first aft support arm 416 b and is operablyconnected in fluid communication with the auxiliary oxidizer tank 458.The second propellant outlet interface 414 b is positioned inboard of,and laterally adjacent to, the base of the second aft support arm 416 cand is operably connected in fluid communication with the auxiliary fueltank 456. A first propellant umbilical 460 a (e.g., a LOX umbilical)extends from the first propellant outlet interface 414 a and has anoutlet 461 a, and a second propellant umbilical 460 b (e.g., a fuelumbilical) extends from the second propellant outlet interface 414 b andhas an outlet 461 b. As described above with reference to FIG. 1E, inoperation the outlet 461 a of the first propellant umbilical 460 a isconfigured to releasably connect to the first propellant inlet interface124 a on the vehicle 100, and the outlet 461 b of the second propellantumbilical 460 b is configured to releasably connect to the secondpropellant inlet interface 124 b on the vehicle 100.

As described in greater detail below, during launch of the vehicle 100,the auxiliary oxidizer tank 458 provides oxidizer (e.g., LOX) to thevehicle oxidizer tank 242 (FIG. 2) for the vehicle main engines 120 a-cvia the first propellant umbilical 460 a, and the auxiliary fuel tank456 provides fuel (e.g., Jet-A) to the vehicle fuel tanks 240 a, b (FIG.2) for the vehicle main engines 120 a-c via the second propellantumbilical 460 b. Additionally, as further described below, in someembodiments the first propellant umbilical 460 a is also configured torecirculate vented oxidizer from the vehicle oxidizer tank 242 back tothe sled 400, and the second propellant umbilical 460 b is alsoconfigured to recirculate vented fuel from the vehicle fuel tanks 240 a,b back to the sled 400. Just prior to separation of the vehicle 100 fromthe sled 400, the valves associated with each of the vehicle propellantinterfaces 124 a, b (FIG. 1E) are closed, the propellant umbilicals 460a, b from the sled 400 are disconnected from the correspondinginterfaces 124 a, b, and the associated interface doors are closed. Bylaunching the vehicle 100 in this way, the sled 400 operates as both a“first stage” and an external propellant tank of the vehicle 100.

In addition to the propellant interfaces 414 a, b, the sled 400 canfurther include an electrical umbilical 426 extending from an electricalinterface 425 positioned just aft of the base of the forward support arm416 a on the top plate 406. As described in greater detail below, thedistal end portion of the electrical umbilical 426 can include one ormore electrical connectors 427 configured to releasably connect tocorresponding receptacles on the vehicle electrical interface 125 (FIG.1E) so that operating commands, power and data can be transmittedbetween the sled 400 (e.g., the sled controller 440) and the vehicle 100while the vehicle is mounted to the sled.

The sled 400 can further include a sled braking system 455 configured toslow the sled 400 to a stop if an anomalous operating condition of oneor more of the sled engines 404 a-c or the main vehicle engines 120 a-cis detected during launch. The sled braking system 455 also slows thesled to a stop at the end of each takeoff run. In some embodiments, thebraking system 455 can be at least generally similar and structure andfunction to the braking system used on the Holloman high speed testtrack at Holloman Air Force Base in New Mexico, U.S.A., which includes awater brake comprising a series of water barriers of calibrated depththat slow and stop the sled. In other embodiments, the braking system455 can include hydraulically-actuated brakes (not shown) that engagethe rails 410 a-c to slow the sled. In other embodiments, the sled 400can include other mechanically, pneumatically, electrically,magnetically, and/or hydraulically actuated forms of braking devicesincluding, for example, reverse thrust rocket engines (also not shown).

The sled 400 can also include a controller 440 having a processor 442and memory 444. The controller 440 can receive inputs 413 and issueoutputs 415 (e.g., commands directing sled operation, includingoperating valves, pumps, actuators, and/or other components). By way ofexample, the inputs 413 can include control signals and commands from,e.g., ground systems, the vehicle 100, etc., data from the vehicle 100,operating parameters such as speed, temperature, etc., engine operatingparameters, propellant parameters, and/or other information). Theoutputs 415 can include commands directing sled operation (includingengine operation, e.g., start, stop and throttle settings, brakingsystem operation, etc.), data transmissions, etc. The processor 442 caninclude one or more logic processing units, such as one or more CPUs,DSPs, ASICs, etc., that are operably connected to controls associatedwith, for example, the engines 404 a-c, the propellant tanks 450, 452,456 and 458, the support struts 416 a-c, the sled braking system 455,etc. The processor 442 can control operation of these sled systems asdescribed herein in accordance with computer-readable instructionsstored on the memory 444.

FIG. 4D is a schematic diagram illustrating propellant distributionbetween the launch sled 400 and the vehicle 100 during launch, inaccordance with an embodiment of the present technology. In thisparticular embodiment, one or more of the sled oxidizer tanks 452 andthe sled fuel tanks 450 provide propellant for both the vehicle engines120 a-c and the corresponding sled engine(s) 404 a-c during the sledrun. In FIG. 4D, the vehicle 100 is mounted to the launch sled 400 inpreparation for launch. For ease of illustration, the oxidizer tanks 452and the fuel tanks 450 on the launch sled 400 are represented by asingle oxidizer tank 452 and a single fuel tank 450, respectively. Inthe illustrated embodiment, the launch sled 400 includes a firstpropellant conduit or line 462 a that provides fuel from the fuel tank450 to the sled engines 404 a-c, and a second propellant line 462 b thatprovides oxidizer from the oxidizer tank 452 to the engines 404 a-c.Although not shown, one or more propellant pumps can be associated witheach of the tanks 450, 452 to drive propellant from the tanks throughthe associated propellant lines 462 a, b, and to the vehicle 100.Additionally, the launch sled 400 further includes a first valve 464 ain fluid communication with the first line 462 a that can direct fuelfrom the fuel tank 450 to the first propellant outlet interface 414 a,and a second valve 464 b in fluid communication with the second line 462b that can direct oxidizer from the oxidizer tank 452 to the secondpropellant outlet interface 414 b. Prior to launch, the first propellantline 460 a is operably connected to the first propellant inlet interface124 a on the vehicle 100, and the second propellant line 460 b isoperably connected to the second propellant inlet interface 124 b.

In the illustrated embodiment, the vehicle 100 includes a firstpropellant line 466 a that provides oxidizer from the vehicle fuel tanks240 a, b to the vehicle main engines 120 a-c, and a second propellantline 466 b that provides fuel from the vehicle oxidizer tank 242 to themain engines 120 a-c. Although not shown, one or more propellant pumpscan be associated with each of the tanks 240 a, b and 242 to drivepropellant from the tanks through the associated propellant lines 466 a,b. Additionally, the vehicle 100 further includes a first valve 468 a influid communication with the first propellant line 466 a, and a secondvalve 468 b in fluid communication with the second propellant line 466b. The first valve 468 a is configured to receive fuel from the firstpropellant inlet interface 124 a and provide the fuel to the engines 120a-c. Similarly, the second valve 468 b is configured to receive oxidizerfrom the second propellant inlet interface 124 b and provide theoxidizer to the engines 120 a-c.

Propellant can be distributed between the launch sled 400 and thevehicle 100 during vehicle takeoff in one embodiment as follows. Priorto takeoff, the valves 464 a, b are opened to provide fuel and oxidizerfrom the fuel and oxidizer tanks 450 and 452, respectively, to the sledengines 404 a-c for ignition. Additionally, the valves 464 a, b alsoprovide fuel and oxidizer to the valves 468 a, b, respectively, on thevehicle 100 via the propellant lines 460 a, b, respectively. Prior toigniting the sled engines 404 a-c, the valves 468 a, b are positioned todirect the fuel and oxidizer from the sled fuel tank 450 and the sledoxidizer tank 452, respectively, to the vehicle main engines 120 a-c forignition. The sled engines 404 a-c and the vehicle engines 120 a-c arethen ignited, and once the sled and vehicle engines come up to launchthrust, the sled 400 is released on the takeoff run down the rails 410a-c (FIG. 4A). Accordingly, the sled fuel tank 450 and sled oxidizertank 452 provide fuel and oxidizer, respectively, to both the sledengines 404 a-c and the vehicle main engines 120 a-c throughout ignitionand the takeoff run of the vehicle 100. Just prior to (or during)separation of the vehicle 100 from the launch sled 400 at the end of thetakeoff run, the valves 464 a, b on the sled 400 are closed, and thepropellant lines 460 a, b are disconnected from the correspondingpropellant inlet interfaces 124 a, b. Just prior to this time, however,the valves 468 a, b on the vehicle 100 are positioned to close off theconnections to the propellant inlet interfaces 124 a, b and insteadenable propellant to flow from the vehicle oxidizer tank 242 and thevehicle fuel tanks 240 a, b to the main engines 120 a-c via the firstline 466 a and the second line 466 b. Thus, the vehicle 100 does notbegin burning its own propellant until just prior to separation from thelaunch sled 400. FIG. 4D illustrates one approach to propellantdistribution between the launch sled 400 and the vehicle 100, and inother embodiments other approaches can be employed, as described in moredetail below.

FIGS. 5A and 5B are side and end views, respectably, of the coupling 454a configured in accordance with embodiments of the present technology.For ease of reference, the coupling 454 a will be referred to herein asthe “coupling 454,” with the understanding that the illustratedembodiment can apply to both of the couplings 454 a and 454 b shown inFIG. 4A. In the illustrated embodiment, the coupling 454 includes a body564 and a plurality of sacrificial inserts 560 a-c. The body 564includes a lug portion 569 that extends upwardly from a pair of legs 570a and 570 b. The body 564 is structurally coupled to the underside ofthe sled 400 by means of a pin or bolt that extends through an opening566 in the lug portion 569. Each of the legs 570 a, b includes aplurality of leg portions 568 a-c that wrap around an upper cap 562 ofthe launch rail 410 to couple the coupling 454 to the rail 410. In oneaspect of this embodiment, each of the sacrificial inserts 560 a-c canbe manufactured from a suitably tough synthetic resin, such aspolytetrafluoroethylene (PTFE). One example of such material is commonlyreferred to as Teflon™. Each of the leg portions 568 a-c has acorresponding sacrificial insert 560 a-c fixedly attached thereto toprovide low friction, load-bearing contact surfaces between the body 564and the launch rail 410. As noted above with reference to FIG. 4A, thecoupling 454 moveably (e.g., slideably) engages the sled 400 with thelaunch rail 410, and enables the sled 400 to move fore and aft on therail 410 while restricting movement perpendicular to the rail (e.g.,side-to-side and up-and-down). In operation, the sacrificial inserts 560a-c are able to withstand the high temperatures and high pressures thatresult from supporting the weight of the vehicle 100 and the sled 400 onthe launch rails 410 during the vehicle takeoff run. After each use, thesacrificial inserts 560 a-c can be inspected to determine condition andeasily replaced if necessary.

In other embodiments, the launch sled 400 can be moveably attached tothe launch rails 410 a-c (FIG. 4A) using other types of suitablecoupling devices. For example, in some embodiments the launch sled 400can be movably coupled to the rails 410 a-c by a magnetic levitation(“Maglev”) system that is at least generally similar in structure andfunction to the Maglev systems found on high speed trains. Such Maglevsystems are described in, for example, U.S. Pat. No. 6,044,770, titled“Integrated High Speed MAGLEV System,” which is incorporated herein byreference in its entirety. The Maglev system can be incorporated intothe rails 410 a-c and can include, for example, a plurality of magnetsto support the sled 400 at relatively high speeds with relativelylittle, if any, friction between the sled 400 and the rails 410 a-c.Additionally, the Maglev system can also include a plurality of magnetsfor propelling the sled 400 down the rails 410 a-c during launch, and/ordecelerating and stopping the sled 400 at the end of the launch run.

In some embodiments, use of magnetic levitation and/or propulsion cansupplement or replace the propulsion provided by the rocket engines 404a-c (FIG. 4A). For example, an electromagnetically levitated rocketpropelled sled 400 can reduce the amount of propellant carried on thesled that is required to reach takeoff velocity and can also reducemaintenance required for the sled system. In some embodiments, the sled400 can include a super conducting electromagnetically levitated(SCMaglev) and propelled sled that will eliminate the need for rocketengines on the sled 400. Such embodiments can significantly reduce thetotal amount of propellant required for takeoff, reduce launch noise,and improve launch turnaround time so that multiple launches can becarried out in a short period of time (e.g., in a single day). In yetother embodiments, the sled 400 can be moveably coupled to the rails 410a-c using a system of rollers. However, if rollers or similar systemsare used, the rollers must be carefully selected to ensure that thebearings do not overheat and fail at the sled speeds required forvehicle take off run.

FIG. 6 is a flow diagram of a method 600 for mounting the vehicle 100 tothe launch sled 400 and preparing the vehicle 100 for launch, inaccordance with embodiments of the present technology. FIG. 7A is afront isometric view of the vehicle 100 mounted to the launch sled 400,and FIGS. 7B-7E are a series of side views illustrating various stagesof the method of FIG. 6 in accordance with embodiments of the presenttechnology. Referring to FIG. 6, the method 600 begins in block 602 withthe vehicle 100 being towed onto the sled 400 on its landing gear 126a-c, as shown in FIG. 7B. In block 604, command and telemetry umbilicalsare connected from the sled 400 to the vehicle 100 and verified, and inblock 606, electrical umbilicals are connected from the sled 400 to thevehicle 100 to provide ground power. In some embodiments, the operationsof blocks 604 and 606 can be performed concurrently by a ground crew byconnecting the electrical umbilical 426 (FIG. 4A) from the sled 400 tothe electrical interface 125 on the vehicle 100 (FIG. 1E), as shown inFIG. 7B.

In block 608, the support arms 416 a-c are rotated upwardly from thesled 400 to an intermediate position to engage the corresponding supportinterfaces 127 a-c on the underside of the vehicle 100 (FIG. 1E), asshown in FIG. 7C. The support arms 416 a-c are releasably coupled to thecorresponding interfaces 127 a-c on the vehicle 100 by suitable “holdand release mechanisms.” Engaging the support arms 416 a-c with thevehicle 100 provides the stability needed to connect the propellantumbilicals 460 a, b from the sled 400 to the vehicle 100, and in block610, the propellant umbilicals 460 a, b are connected from the sled 400to the corresponding propellant interfaces 124 a, b on the underside ofthe vehicle 100 (FIG. 1E), as also shown in FIG. 7C. (It should be notedthat the propellant umbilicals 460 a, b are hidden behind the first aftsupport arm 416 b in FIG. 7C.)

Once the propellant umbilicals 460 a, b have been connected to thevehicle 100, their connections are verified, and in block 612, thesupport arms 416 a-c are further rotated upwardly enough to raise thevehicle 100 off of its landing gear 126 a-c (see arrows in FIG. 7D). Inblock 614, the landing gear 126 a-c retract into their respective gearbays on the vehicle 100, as shown in FIG. 7D, and the gear doors 129 a,care closed and sealed. In block 616, all of the umbilical connectionsbetween the launch sled 400 and the vehicle 100 (e.g., the connectionsbetween the electrical umbilical 426 and the propellant umbilicals 460a, b and the vehicle 100) are verified to ensure that, for example, theelectrical connections have sufficient integrity, and that thepropellant connections are properly sealed. Once this has been done, inblock 618, the support arms 416 a-c are rotated further aft to raise thevehicle 100 to its release angle of attack (AOA) as shown in FIG. 7E. Inother embodiments, the landing gear 126 a-c can remain down until thevehicle is raised to its release AOA as shown in FIG. 7E, and can thenbe retracted into their respective gear bays. In one aspect of theseembodiments, it should be noted that because the electrical umbilical426 and propellant umbilicals 460 a, b are positioned adjacent to therespective support arms 416 a-c and are approximately the same length asthe respective support arms 416 a-c, the length of the umbilicals doesnot change substantially when the support arms 416 a-c are rotated fromthe position shown in FIG. 7D to the position shown in FIG. 7E. Thisavoids the need to provide for a substantial change in length of theumbilicals 426 and 460 a, b during rotation of the vehicle 100.

FIG. 8A is a partially schematic side view illustrating the connectionsbetween the vehicle 100 and the forward support arm 416 a and theelectrical umbilical 426, and FIG. 8B is a partially schematic rear viewillustrating the connections between the vehicle 100 and the rearsupport arms 416 b, c and the propellant umbilicals 460 a, b, inaccordance with embodiments of the present technology. Referring firstto FIG. 8A, as noted above, the distal end portion of each of thesupport arms 416 a-c can be releasably attached to the correspondingsupport arm interface 127 a-c by a hold and release mechanism 810 untilcommanded to release. Although the hold and release mechanisms 810securely attach the vehicle 100 to the support arms 416 a-c, asdescribed in greater detail below with reference to FIGS. 9A-12D, thehold and release mechanisms 810 enable the support arms 416 a-c torotate relative to the vehicle 100 to move the vehicle 100 relative tothe sled 400 as described above with reference to FIGS. 7B-7D. The holdand release mechanisms 810 can include, for example, one or moremechanical hold-down arms connected to linkages that are pneumatically,hydraulically, and/or electrically controlled to hold an engagementfeature (e.g., an engagement surface) on the corresponding interface 127a-c until commanded to release in response to a control command. In someembodiments, the hold and release mechanism 810 can be carried on thedistal end portion of each support arm 416 a-c to reduce vehicle weight.In other embodiments, all or a portion of the hold and release mechanism810 can be incorporated into the vehicle interface 127 a-c. Hold andrelease mechanisms are known in the art, and in other embodiments, othersuitable hold and release systems can be used to attach the support arms416 a-c to the corresponding interfaces 127 a-c until commanded torelease. As also shown in FIG. 8A, each of the support arm interfaces127 a-c can include a corresponding door 802 that automatically closesonce the support arm 416 a-c has been released, to thereby close off theinterface and protect it from detrimental aerothermal heating, etc.

In some embodiments, the electrical umbilical 426 can be attached to thecorresponding electrical interface 125 on the vehicle 100 using anynumber of suitable connector/receptacle mechanisms known in the art tomaintain the electrical connections between the sled 400 and the vehicle100 until the electrical umbilical 426 is retracted (via, e.g., alanyard connected to the sled 400) for vehicle separation and liftoff.As schematically illustrated in FIG. 8A, in some embodiments electricalpower and commands are transmitted from the vehicle 100 to the sled 400via the electrical umbilical 426, and data/feedback can be transmittedfrom the sled 400 to the vehicle 100 via the electrical umbilical 426.In other embodiments, power, commands, data, and/or other electricalinformation and signals can be transmitted in different directions viathe electrical umbilical 426. As with the support arms 416 a-c, theelectrical interface 125 and both of the propellant interfaces 124 a, bcan also include a door 804 that automatically closes upon retraction ofthe corresponding umbilical.

Referring next to FIG. 8B, in the illustrated embodiment oxidizer (e.g.,subcooled LOX) flows from the sled 400 to the vehicle 100 via the firstpropellant umbilical 460 a. In one aspect of this embodiment, the firstpropellant umbilical 460 a can be further configured to enable oxidizerthat is vented from the vehicle oxidizer tank 242 to flow back to thesled 400. For example, in some embodiments the first propellantumbilical 460 a can include two separate conduits: one that flowsoxidizer (e.g., subcooled LOX) to the vehicle tank 242, and one thatflows vented oxidizer from the vehicle tank 242 back to the sled 400 forre-cooling and re-densifying. As described in greater detail below, byrecirculating vented oxidizer back to the launch sled 400 in theforegoing manner, the vehicle oxidizer tank 242 can be maintained at afull level. Similarly, the second propellant umbilical 460 b can alsoinclude two separate conduits: one configured to flow fuel (e.g., Jet-A)from the launch sled 400 to the vehicle 100, and another to recirculatevented fuel from the vehicle tanks 240 a, b back to the sled so that thevehicle tanks 240 a, b are maintained at a full fuel level. The outlets461 a, b of the propellant umbilicals 460 a, b can be releasablyattached to the corresponding propellant interfaces 124 a, b usingsuitable propellant conduit couplings known in the art. Such couplingscan include, for example, pneumatically actuated clamps that maintainthe sealed connection between the umbilical outlets 461 a, b and thecorresponding interfaces 124 a, b until commanded to release. In theillustrated embodiment, the second propellant line 460 b is alsoconfigured to provide one or more electrical umbilical connectionsbetween the launch sled and the vehicle 100. Although it isschematically illustrated that the aforementioned electrical umbilicalsextend through the second propellant umbilical 460 b, it will beunderstood that in other embodiments the electrical connections providedat the second propellant umbilical 460 b can be positioned outside ofthe propellant line 460 b or otherwise in close proximity to theumbilical 460 b.

FIGS. 9A and 9B are side and rear views, respectively, of a support armhold and release mechanism 810 in a first stage of operation, and FIG.9C is a side view of the hold and release mechanism 810 in a secondstage of operation, in accordance with embodiments of the presenttechnology. Referring first to FIGS. 9A and 9B, each of the support arms416 a-c on the sled 400 carries a hold and release mechanism 810 on adistal end portion thereof. In some embodiments, each of the hold andrelease mechanisms 810 includes a fitting 980 (e.g., a “backstop”fitting) that is pivotally attached to the distal end portion of thecorresponding support arm 416 a-c by means of a pivot shaft 982.Additionally, each of the hold and release mechanisms 810 includes ahold-down arm 984 that is pivotally coupled to the backstop fitting 980by means of a pivot shaft 985. The distal end portion of the hold-downarm 984 includes a recessed surface 988 that in operation clamps anattachment fitting 970 of the corresponding vehicle interface 127 a-c tothe backstop fitting 980.

FIGS. 10A and 10B are front and side cross-sectional views,respectively, of the attachment fitting 970 configured in accordancewith embodiments of the present technology. Referring to FIGS. 10A and10B together, in some embodiments the attachment fitting 970 can be a“U”-shaped fitting having a curved (e.g., rounded convex) clampingsurface 1072. As shown in FIG. 9A, the recessed surface 988 on thedistal end portion of the hold-down arm 984 has a complementary shape(e.g., rounded concave) to clamp against the surface 1072 of theattachment fitting 970.

FIGS. 11A and 11B are front and side cross-sectional views,respectively, of the backstop fitting 980 configured in accordance withembodiments of the present technology. As these views illustrate, insome embodiments the backstop fitting 980 can include a pocket 989(e.g., a rectangular-shaped pocket) having side walls 990 a, b and aback wall 990 c. The pocket 989 is shaped and sized to receive andretain the attachment fitting 970 against downward and aft movement, yetenable the attachment fitting 970 to move forward and upwardly out ofthe backstop fitting 980 when released by the hold-down arm 984, asdescribed in greater detail below.

As described above, the support arms 416 a-c are configured to rotate asshown in FIGS. 7D and 7E after attachment to the corresponding vehicleinterfaces 127 a-c. To accommodate this rotation, the hold and releasemechanism 810 includes extensible actuators 992 a, b, each having adistal end portion attached to the backstop fitting 980. The actuators992 a, b are configured to extend as the corresponding control arms 416a-c rotate aft, thereby causing the corresponding backstop fitting 980to rotate about the pivot shaft 982 and accommodate rotation of thesupport arm 416 a-c relative to the corresponding vehicle attachmentfitting 970. In some embodiments, the actuators 992 a, b can behydraulic actuators, pneumatic actuators, electro-mechanical actuators,or other suitable types of actuators, including other nonlinearactuators known in the art.

As shown in FIG. 9C, when the hold and release mechanism 810 iscommanded to release, the hold-down arm 984 rotates away from theattachment fitting 970, thereby enabling the fitting 970 to moveupwardly and forwardly out of the backstop fitting 980 as the vehicle100 (FIG. 1A) separates away from the sled 400 (FIG. 4A). Movement ofthe hold-down arm 984 can be controlled by various suitable drivesystems known in the art. For example, in some embodiments the hold-downarm 984 can be held in the “hold” position (FIG. 9A) and then moved tothe “release” position (FIG. 9C) by a linkage that is hydraulically orpneumatically actuated. In other embodiments, movement of the hold-downarm 984 can be controlled by an electro-mechanical system.

The hold and release mechanism 810 described above with reference toFIGS. 9A-11B is one example of a suitable hold and release system thatcan be used with embodiments of the present technology. In otherembodiments, other suitable hold and release mechanisms can be usedwithout departing from the present disclosure. For example, FIGS. 12Aand 12B are side and rear views, respectively, of a hold and releasemechanism 1210 in a first stage of operation, and FIGS. 12C and 12D arecorresponding side and rear views, respectively, of the hold and releasemechanism 1210 in a second stage of operation, in accordance withembodiments of the present technology. Referring first to FIGS. 12A and12B, the hold and release mechanism 1210 can be referred to as a“clamshell” type hold and release mechanism that utilizes two hold-downmembers 1274 a, b to securely attach an end fitting 1270 at each of thevehicle interfaces 127 a-c to the corresponding support arm 416 a-c. Inthe illustrated embodiment, the hold-down members 1274 a, b are arcuatering segments (or “clamshells”) having an inner diameter configured toencircle a cylindrical crossbar 1272 that extends horizontally on theend fitting 1270. The proximal end portions of the first and secondhold-down members 1274 a, b are pivotally coupled to a base 1290 bymeans of corresponding pivot pins 1276 a, b. The distal end portions ofthe first and second hold-down members 1274 a, b are releasably securedto the base 1290 in the closed position by a lockpin 1278 that extendsthrough a bore in the base 1290 and through corresponding holes (notshown) in the distal end portions.

Turning next to FIGS. 12C and 12D, when the hold and release mechanism1210 is commanded to release, the lockpin 1278 is retracted from thedistal end portions of the first and second hold-down members 1274 a, band the hold-down members 1274 a, b are rotated aft about thecorresponding pivot pins 1276 a, b. Movement of the hold-down members1274 a, b in the foregoing manner releases the crossbar 1272 and enablesthe attachment fitting 1270 to move forwardly and upwardly away from thehold and release mechanism 1210 as the vehicle 100 separates from thesled 400. The hold and release mechanisms described above are providedby way of example only. Accordingly, those of ordinary skill in the artwill appreciate that other suitable hold and release mechanisms can beused, consistent with the present disclosure.

FIG. 13 is a schematic diagram illustrating a ground based propellantmanagement system 1300 connected to the vehicle 100 and the sled 400 inaccordance with embodiments of the present technology. In someembodiments, the propellant management system 1300 includes an oxidizermanagement system 1302 (e.g., a LOX management system) and a fuelmanagement system 1304 which are operably connected to a controller1340. Those of ordinary skill in the art will understand that theoxidizer management system 1302 and the fuel management system 1304include electronically controlled propellant pumps, valves, actuators,and associated propellant conduits, etc. configured to direct thepropellant flows in the manner described herein. The propellantmanagement system controller 1340 includes one or more processors thatcontrol operations and functions of the oxidizer management system 1302,the fuel management system 1304, and other components of the propellantmanagement system 1300 in accordance with computer-readable instructionsstored on a suitable memory. It will be understood that the controller1340 can include any logic processing unit, such as one or more CPUs,DSPs, APUs, etc. The propellant management system 1300 further includesa heat exchanger 1306 and a liquid nitrogen (LN2) supply 1308. The LN2supply 1308 is configured to circulate LN2 through the heat exchanger1306 to increase the density of LOX flowing through the heat exchanger1306 by lowering its temperature from the normal boiling point (NBP) ofLOX (i.e., −182.96 degrees C.) to the NBP of LN2 (i.e., −195.79 degreesC.). It will be appreciated that although the system of FIG. 13 isdescribed in the context of a LOX management system, such system couldbe used with other types of oxidizer.

In the illustrated embodiment, the propellant management system 1300includes a LOX feed line 1310 that can be releasably connected in fluidcommunication to the sled LOX tanks 452, and a fuel feed line 1312 thatcan be releasably connected in fluid communication to the sled fueltanks 450. Although the sled 400 includes three LOX tanks 452 and threefuel tanks 450 in some embodiments, the LOX tanks 452 and the fuel tanks450 are depicted as a single LOX tank and a single fuel tank,respectively, in FIG. 13 for ease of illustration. The propellantmanagement system 1300 further includes a LOX return line 1318 and afuel return line 1330. The LOX return line 1318 can be releasablyconnected in fluid communication to the sled LOX tanks 452, theauxiliary LOX tank 458, and the vehicle LOX tank 242 to enable ventedLOX to recirculate through the heat exchanger 1306. The fuel return line1330 can be releasably connected to the sled 400 in fluid communicationwith the vehicle fuel tank 240.

FIG. 14 is a flow diagram of a routine 1400 for loading propellants ontothe vehicle 100 and the sled 400 in accordance with embodiments of thepresent technology. All or portions of the routine 1400 can be performedby ground control computers, the propellant management system controller1340, and/or computers onboard the vehicle 100 in accordance withcomputer readable instructions stored on memory. Referring to FIG. 14with reference to FIG. 13, in block 1402, once the vehicle 100 is at itsliftoff angle of attack, NBP LOX is circulated through the LN2 heatexchanger 1306. This increases the density of the LOX by lowering itstemperature from LOX NBP (i.e., −182.96 degrees C.), or from about LOXNBP, to LN2 NBP (i.e., −195.79 degrees C.), or to about LN2 NBP. Inblock 1404, the LOX management system 1302 transfers the densified LOXfrom the heat exchanger 1306 to the sled LOX tanks 452 and the auxiliaryLOX tank 458 via the LOX feed line 1310. It will be understood that, insome embodiments the propellant management system 1300 and/or the sled400 can include one or more propellant pumps and/or associated valvesoperably connected to the LOX feed line 1310 to facilitate and controlthe transfer of LOX from the heat exchanger 1306 to the sled LOX tanks452 and the auxiliary LOX tank 458. As the LOX in the tanks 452 and 458warms to its NBP, it expands and is vented from the tanks and returns tothe heat exchanger 1306 via the recirculation line 1318 so that the LOXcan be re-densified. The heat exchanger 1306 maintains the LOX at thetarget temperature of, e.g., −195.79 degrees C., or about −196 degreesC., and a target pressure of, e.g., about 17 psia. In other embodiments,the heat exchanger 1306 can maintain the LOX at other targettemperatures and/or target pressures. For example, in some embodimentsthe heat exchanger 1306 (and/or other components of the propellantmanagement system 1300) can maintain the LOX at a target pressure lessthan 17 psia, or greater than 17 psia, such as about 20 psia, or about40 psia.

In block 1406, once the temperature and pressure of the LOX in the sledtanks 452 and 458 have stabilized at the target temperature andpressure, densified LOX is loaded into the vehicle LOX tank 242 from thesled auxiliary LOX tank 458 via the first propellant umbilical 460 a. Itwill be understood that, in some embodiments the sled 400 and/or thevehicle 100 can include one or more propellant pumps and/or associatedvalves operably connected in fluid communication with the firstpropellant umbilical 460 a to facilitate and control the transfer of LOXfrom the sled auxiliary LOX tank 458 into the vehicle LOX tank 242. Asthe LOX in the vehicle LOX tank 242 warms and expands, it is vented fromthe LOX tank 242 through a recirculation umbilical 1320 back to the sled400, and then back to the heat exchanger 1306 via the recirculation line1318. Although the recirculation umbilical 1320 is depicted as beingseparate from the first propellant umbilical 460 a for ease ofillustration, it will be understood from the description of the firstpropellant umbilical 460 a provided above with reference to FIG. 8Bthat, in some embodiments, the LOX recirculation umbilical 1320 can beconjoined or otherwise combined with the first propellant umbilical 460a.

In block 1408, the LOX flowing back to the heat exchanger 1306 is cooledand densified as described above and can be recirculated back to thesled LOX tanks 452, the sled auxiliary LOX tank 458, and the vehicle LOXtank 242, as needed to maintain the tanks in full, or at least nearlyfull, conditions. In block 1410, once the LOX in the sled and vehicletanks has stabilized, the fuel management system 1304 begins circulatingfuel through a fuel heat exchanger (not shown) until it reaches itstarget temperature and pressure (e.g., about 60 degrees F. and about 40psia). In other embodiments, the fuel heat exchanger can maintain thefuel at other target temperatures and/or target pressures. For example,in some embodiments the fuel heat exchanger (and/or other components ofthe propellant management system 1300) can maintain the fuel at a targetpressure less than 40 psia, such as about 17 psia, or about 20 psia, orgreater than 40 psia. In block 1412, the fuel is then loaded into thesled fuel tanks 450 and the sled auxiliary fuel tank 456 via the fuelfeed line 1312 and circulated to maintain the fuel at its targettemperature and pressure. It will be understood that, in someembodiments the propellant management system 1300 and/or the sled 400can include one or more propellant pumps and/or associated valvesoperably connected to the fuel feed line 1312 to facilitate and controlthe transfer of fuel from the fuel management system 1304 to the sledfuel tanks 450 and the sled auxiliary fuel tank 456. In block 1414, fuelthen flows from the auxiliary fuel tank 456 on the sled 400 to thevehicle fuel tank 240 via the second propellant umbilical 460 b. In someembodiments, the sled 400 and/or the vehicle 100 can include one or morepropellant pumps and/or associated valves operably connected in fluidcommunication with the second propellant umbilical 460 b to facilitateand control the transfer of fuel from the sled auxiliary fuel tank 456into the vehicle fuel tank 240. The fuel circulates from the vehiclefuel tank 240 back to the sled 400 via a recirculation line 1326.Although the recirculation line 1326 is illustrated as being separatefrom the second propellant umbilical 460 b for ease of illustration, asdescribed above with reference to FIG. 8B, in some embodiments the fuelrecirculation path from the vehicle 100 to the sled 400 can be through aconduit that is carried by the second propellant umbilical 460 b. Therecirculated fuel then flows from the sled 400 back to the fuelmanagement system 1304 via the fuel recirculation line 1330 to maintainthe fuel at the target temperature and pressure. Circulation of the fuelthrough the tanks 450, 456 and 240 and the fuel management system heatexchanger as described above ensures that the fuel maintains its targettemperature and pressure throughout the system.

FIG. 15 is a schematic diagram of a suitable control system architecturefor performing all or a portion of the routines described above when thevehicle 100 is connected to the propellant management system 1300, inaccordance with embodiments of the present technology. In theillustrated embodiment, ground control 1500 (e.g., one or morecontrollers and/or other processing devices executing computer readableinstructions and/or responding to user inputs) controls the propellantmanagement system 1300, ensuring that the temperature, pressure and flowrates of the propellants remain constant, or at least approximatelyconstant, as the propellants flow through the sled 400 and the vehicle100. The propellant management system 1300 provides the status of thepropellants and other relevant data about the propellant managementsystem (e.g., data feedback) to ground control 1500 and the vehicle 100.Ground control 1500 also controls operations of the vehicle 100 at thistime and, through the vehicle 100, is able to control the sled 400. Inother embodiments, ground control 1500 can control operation of the sled400 directly. The sled provides data feedback to the vehicle 100, whichin turn provides data feedback from both the sled and the vehicle 100 toground control 1500. Ground control and the vehicle 100 remain inconstant communications (e.g., verbal communications from the crew ofthe vehicle 100 to ground control personnel) throughout the entireprocess with ground control 1500 providing updates to the crew of thevehicle 100 and the crew providing information to ground control, whennecessary. In other embodiments, the communication, control, and/orfeedback paths between ground control 1500, the vehicle 100 and/or thesled 400 can differ from those described above. As further illustratedin FIG. 15, and as described in detail above, the propellant managementsystem 1300 provides propellants to both the sled and the vehicle andmaintains the propellants at target temperatures and pressures viarecirculation. It will be understood by those of ordinary skill in theart that the communications, controls, and data feedback providedbetween the ground control 1500, the propellant management system 1300,the sled 400, and the vehicle 100 will generally be implemented viawired and/or wireless connections providing digital communications ofinformation, data, control signals, etc. between the respective systemcontrollers.

FIG. 16A is a flow diagram of a routine 1600 a for operating the vehicle100 and the launch sled 400 during a takeoff run; FIG. 16B is a flowdiagram of a routine 1600 b for confirming safe liftoff conditions priorto release of the vehicle 100 from the sled 400; FIG. 16C is a flowdiagram of a routine 1600 c for liftoff of the vehicle 100 from the sled400; and FIG. 16D is a flow diagram of a routine 1600 d for abortingliftoff of the vehicle 100 from the sled 400, in accordance withembodiments of the present technology. FIGS. 17A-17D are a series ofschematic diagrams illustrating operation of the vehicle 100 and thelaunch sled 400 at various stages of the launch sequences described inFIGS. 16A-16D, in accordance with embodiments of the present technology.Referring first to the routine 1600 a of FIG. 16A, in block 1602, onceall of the pre-flight checkouts for the vehicle 100 have been completedand the vehicle is ready for takeoff, ground control stops propellantflow from the propellant management system 1300 (FIG. 13) to the sled400 and then disconnects the propellant management system 1300 from thesled 400, as shown in FIG. 17A. In block 1604, the sled propellantsystem continues to maintain the sled and vehicle propellant conditionsuntil engine ignition by recirculation of LOX and fuel from the vehicleLOX tank 242 and the vehicle fuel tank 240, respectively, to the sledLOX tanks 452 and 458, and the sled fuel tanks 450 and 456,respectively, as also shown in FIG. 17A.

In block 1606, ground control commands the sled and vehicle engines 404a-c and 120 a-c, respectively, to ignite and throttle up to 100 percent,as shown in FIG. 17B. In block 1608, all the engines 120 a-c on thevehicle 100 are then set to “tripped flow” for optimal, or at least nearoptimal, low altitude performance. In some embodiments, setting theengines 120 a-c to tripped flow changes the area ratio of the enginesfrom, e.g., 60:1 to, e.g., 33:1. Together, the vehicle engines 120 a-cand the sled engines 404 a-c provide thrust to propel the sled 400 andthe vehicle 100 down the launch rails 410 a-c (FIG. 4A) for takeoff. Inblock 1610, throughout the entire sled run, the sled LOX and fuel tanks452 and 450, respectively, feed propellants to the sled engines 404 a-c,while the auxiliary LOX and fuel tanks 458 and 456, respectively, refuelthe vehicle LOX and fuel tanks 242 and 240, respectively. As describedabove and elsewhere herein, refueling of the vehicle propellant tanks inthe foregoing manner enables the vehicle 100 to lift off from the sled400 with full, or at least approximately full, propellant tanks, therebyeliminating, or at least greatly reducing, the liftoff dry weightpenalty associated with conventional space launch vehicles.

In some embodiments, the vehicle 100 is in the takeoff angle of attackposition shown in FIG. 7E during the takeoff run. In other embodiments,instead of being positioned at the takeoff angle of attack during theentire takeoff run, the vehicle 100 can begin the takeoff run at a lowerangle of attack (such as the position shown in FIG. 7D) to reduce dragand increase acceleration, and then the support arms 416 a-c can rotateaft to increase the angle of attack of the vehicle 100 just prior toliftoff and separation from the sled 400.

If any anomalies are detected during any part of the sled run afterblock 1610, the routine proceeds to block 1616 and initiates a refusedtakeoff sequence. Alternatively, if no anomalies are detected, theroutine proceeds to block 1612 and, at or just before reaching takeoffspeed, the propellant flow from the sled 400 to the vehicle 100 isstopped, the first and second propellant umbilicals 460 a and 460 b aredisconnected and retracted from the vehicle 100, and the propellantdoors 804 (FIG. 8A) on the vehicle 100 are closed, as depicted by FIG.17C. From this point forward, the vehicle engines 120 a-c are usingpropellant solely from the vehicle LOX and fuel tanks 252 and 240,respectively. If anomalies are detected during or after the separationprocess of block 1612, the routine proceeds to block 1616 and initiatesthe refused takeoff sequence. If not, then the routine proceeds to block1614 and the vehicle and sled systems (e.g., the respective controllers)check their respective operating systems to confirm that safe liftoffconditions have been met.

Referring next to FIG. 16B, the routine 1600 b starts at block 1614 fromthe routine 1600 a above. In block 1614, just prior to vehicleseparation from the sled 400, the vehicle 100 and the sled 400 perform anumber of system checks to confirm that safe liftoff conditions are met.The system checks can be run concurrently, and can include, for example,confirming that all engines on the vehicle 100 and the sled 400 areoperating nominally (block 1620), all flight critical systems areoperating nominally (block 1622), and the lift on the vehicle is equalto or greater than 1.1 times the gross takeoff weight of the vehicle(block 1624). If one or more of the system checks fail or areunsuccessful, the routine proceeds to block 1626 and executes therefused takeoff sequence. Conversely, if all of the system checks aresuccessful, the routine proceeds to block 1628 and the hold and releasemechanisms on the sled support arms 416 a-c are commanded to release thevehicle 100 as depicted in FIG. 17D.

Referring next to FIG. 16C, the routine 1600 c starts at block 1628 fromthe routine 1600 b above. In block 1628, once the hold and releasemechanisms on the sled support arms have been commanded to release thevehicle 100, the routine proceeds to block 1630 and the vehicle 100separates and lifts off from the sled 400. In block 1632, theelectrical/data umbilicals (e.g., the electrical umbilical 426 (FIG.4A)) are detached from the vehicle interface 125 (FIG. 1E) and retracted(e.g., retracted onto the forward support arm 416 a (FIG. 4A)), and inblock 1634 all remaining doors (e.g., the support arm doors 802 and theelectrical/propellant umbilical doors 804 (FIG. 8A)) are closed. At thistime, the vehicle controller 140 can confirm that all of the electricaland propellant umbilicals have been disconnected from the vehicle 100and all of the associated doors on the vehicle 100 have been closed. Inblock 1636, the sled engines 404 a-c are shut down and in block 1638 thesled slows down and comes to a stop when it encounters the waterbarriers at the end of the sled run.

Turning next to FIG. 16D, as noted above FIG. 16D is a flow diagram of aroutine 1600 d for executing a refused takeoff sequence in accordancewith embodiments of the present technology. In block 1640, the routinebegins when anomalies are detected at any point during the sled run. Inblock 1642, all engines on the sled 400 and the vehicle 100 areimmediately commanded to shut down, and in block 1644 all sled andvehicle systems are commanded to safe conditions. Such safe conditionscan include, for example, opening tank vents on the LOX tanks on boththe sled 400 and the vehicle 100. In block 1646, the sled 400 with thevehicle 100 still mounted thereon slows and safely brakes to a stop whenthe sled 400 encounters the water barriers at the end of the sled run.

FIG. 18A is a schematic diagram of a suitable control systemarchitecture 1800 a for performing all or a portion of the routines 1600a, 1600 b, and 1600 d described above after the vehicle 100 and the sled400 have detached from the propellant management system 1300, inaccordance with embodiments of the present technology, and FIG. 18B is aschematic diagram of a suitable control system architecture 1800 b forperforming all or a portion of the routine 1600 c after the vehicle 100has detached from the sled 400, in accordance with embodiments of thepresent technology. Referring first to FIG. 18A, when the vehicle 100and the sled 400 have disconnected from the propellant management system1300, the vehicle 100 is in control of the sled 400 and provides bothelectrical power and control actuation commands to the sled 400. Thesled 400 provides data feedback on the commanded controls from thevehicle 100 and on the status of the propellant as it continues flowingfrom the sled 400 into the vehicle 100. The vehicle 100 provides datafeedback on all systems in both the sled 400 and the vehicle 100 toground control 1500. In some embodiments, ground control 1500 is nolonger able to control the vehicle 100 at this point, but both maintainconstant communications (e.g., wireless communications) throughout theentire process, providing information to the other when necessary. Inother embodiments, ground control 1500 can provide all, or a portion, ofthe control inputs for the vehicle 100 after separation from the sled400. Accordingly, it will be understood that ground control 1500includes suitable communications systems for wireless transmission ofcommunications, command signals, and telemetry to and from the vehicle100.

Turning next to FIG. 18B, once the vehicle 100 has disconnected from thesled 400, the vehicle 100 is able to provide data feedback from itsonboard systems to ground control 1500, but ground control 1500 is notable to directly control the vehicle 100. Instead, vehicle guidance,navigation, control, systems management, etc. is performed by thevehicle controller 140 in accordance with flight control software andtrajectory optimization code tasks, as described in greater detailbelow. As noted above, however, in other embodiments ground control 1500can provide all or a portion of the control commands and/or inputnecessary for vehicle guidance, navigation, control, and/or systemsmanagement. Throughout the flight, ground control 1500 and the vehicle100 can maintain constant communication, providing information to theother when necessary.

FIG. 19 is a partially schematic diagram illustrating various phases ina flight sequence of the vehicle 100 in accordance with some embodimentsof the present technology. In a takeoff phase 1901, the sled 400 ismounted to the launch rails 410 a-c as described above with reference toFIG. 4A, and the vehicle 100 is mounted to the sled 400 as describedabove with reference to FIG. 7E. On a typical flight, the vehicle 100may carry a payload of about 5,000-7,500 lbs. and a crew of five. Priorto takeoff, the vehicle engines 120 a-c (FIG. 1A) and the sled enginesfor 404 a-c are ignited and brought up to full thrust using propellantfrom the sled 400 as described above. If needed, the sled 400 can beheld in place on the rails 410 a-c as the engines are brought up to fullthrust using a sled braking system or a hold and release mechanism.During the takeoff run, the additional thrust from the sled 400 booststhe vehicle acceleration, and the use of sled propellants instead ofvehicle propellants enables the vehicle 100 to take off fully loadedwith propellant. At or near takeoff speed, the electrical and propellantumbilicals disconnect from the vehicle 100 and the vehicle 100 releasesfrom the support arms 416 a-c and enters a pull up phase 1902.

In some embodiments, the vehicle 100 can achieve a takeoff speed of fromabout 400 mph to about 500 mph, or about 436 mph (0.7 Mach), in about 20seconds after traveling down the rails 410 a-c a distance of from about4,500 ft. to about 6,000 ft., or about 5,182 ft. In some embodiments,the vehicle 100 and crew may experience relatively low dynamic forcesduring takeoff, with maximum accelerations ranging from about 1 g toabout 2 g's, or about 1.42 g's. The foregoing launch parameters areillustrative of some embodiments of the present technology. In otherembodiments, the vehicle 100 can achieve different takeoff speeds indifferent takeoff run distances, and resulting in different maximumacceleration levels, without departing from the present disclosure.

As described above with reference to FIG. 4A, the sled 400 can include abraking system 455. In some embodiments, the braking system 455 canenable the sled 400 to decelerate, with the vehicle 100 attached to thesled 400, from the takeoff speed to zero in a distance of from about3,000 ft. to about 3,500 ft., or about 3,200 ft., resulting in a maximumdeceleration of from about 1 g to about 3 g's, or about 2 g's. This sledbraking feature can enable the sled 400 to perform a refused takeoff atany point prior to takeoff if necessary for mission safety.

After the vehicle 100 lifts off from the sled 400, the vehicle flightpath is controlled by operation of the aerodynamic control surfacesdescribed above with reference to FIG. 1A and/or gimballing the engines120 a-c. The pull up phase 1902 can be relatively gentle and permit alow velocity turn to a target azimuth that provides orbital inclinationflexibility. In some embodiments, the maximum accelerations the vehicle100 experiences during the pull up phase 1902 can be from about 1.2 g'sto about 2 g's, or about 1.7 g's. After the pull up phase 1902, thevehicle 100 enters an ascent phase 1903 in which the vehicle may belimited to maximum accelerations of, for example, about 3 g's tomaintain crew/passenger comfort. Vehicle directional control during allor a portion of the ascent phase can be provided, or at leastsupplemented, by engine thrust vector control and/or engine gimbaling.During the ascent phase 1903, the vehicle can experience a maximumdynamic pressure (max Q) of from about 1,000 psf to about 1,100 psf, orabout 1,080 psf. In some embodiments, at the end of the ascent phase1903 the vehicle 100 will be traveling at a speed of from about 15,000mph to about 20,000 mph, or about 17,560 mph or more, and the vehiclewill be at an altitude of from about 275,000 ft. to about 325,000 ft.,or about 300,000 ft. or more.

Turning next to FIG. 20, this Figure is a flow diagram of a routine 2000for performing an ascent sequence of the vehicle 100 in accordance withembodiments of the present technology. Upon separation from the sled 400in block 2002, the vehicle 100 maintains both the launch azimuth and apreset rate of climb until the aerodynamic effects from separation aredampened. In block 2004, the vehicle continues the ascent and banks tothe target azimuth as commanded by the vehicle controller 140 (FIG. 1A)in accordance with flight control software executing a trajectoryoptimization code. In block 2006, when the optimal altitude has beenreached for changing the area ratio of the nozzles of the engines 120a-c (e.g., typically about 32,000 ft.), the tripped flow is turned offand the engine area ratio is increased from, e.g., about 33:1 to, e.g.,about 60:1. In some embodiments, an area ratio of 60:1, or at leastapproximately 60:1, can provide optimal, or near optimal, high altitudeflight performance of the vehicle engines 120 a-c.

In block 2008, all three of the engines 120 a-c maintain 100 percentthrottle until, in block 2010, a preset axial acceleration limit isreached. In some embodiments, the preset axial acceleration limit can be3 g's. In other embodiments, other preset axial acceleration limits canbe used that are higher or lower than 3 g's. Once the axial accelerationlimit has been reached, all the engines 120 a-c begin throttling down tomaintain the axial acceleration at or below the limit. When the lowerthrottle limit is reached on the engines 120 a-c (typically about 50percent throttle), the outboard engines 120 a and 120 c are shut down,and the center engine 120 b (FIG. 1A) is throttled up to 100 percent. Inblock 2014, when the preset axial acceleration limit is again reached,the center engine 120 b begins throttling down until it reaches itslower throttle limit (e.g., about 50 percent). In block 2020, uponvehicle insertion into the target orbit, the vehicle center engine 120 bshuts down. Final adjustments to the orbit of the vehicle 100 can thenbe performed using the OMS engines 122 a, b (FIG. 1A).

As noted above, all or a portion of the vehicle trajectory and controlduring the ascent sequence of FIG. 20, and/or other portions of theflight of the vehicle 100, can be controlled by the vehicle controller140 in response to execution by the processer 142 of computer-readableinstructions stored on non-volatile memory 144 (FIG. 1A). Thecomputer-readable instructions executed by the processor 142 can includeflight control software and trajectory optimization code. In someembodiments, the trajectory optimization code can include tasks forflight profile optimization and engine management. For example, in someembodiments the trajectory optimization code tasks can includeoptimizing the vehicle angle of attack profile and the vehicle bankangle profile to achieve the desired ascent trajectory. Additionally, insome embodiments the code tasks can also include optimizing enginecontrol to determine, for example, when to throttle down the engines,shut off the outboard engines, change the engine area ratios, etc. Thetrajectory optimization code tasks can also include maintaining flightpath constraints. For example, in some embodiments this can includemaintaining normal acceleration (i.e., acceleration along an axisperpendicular to the longitudinal axis of the vehicle and extending fromthe bottom of the vehicle to the top of the vehicle) at less than orequal to 2.5 g's, maintaining axial acceleration at less than or equalto 3 g's, and maintaining vehicle dynamic pressure at less than or equalto 1,200 psf. The foregoing flight path constraints are examples of somesuitable constraints for some embodiments of the present technology.Accordingly, other embodiments can utilize other flight pathconstraints. In addition to the foregoing, the trajectory optimizationcode tasks can further include targeting the desired final conditionsupon insertion into the target orbit. Such conditions can include, forexample, maximizing the final vehicle weight and achieving presetorbital targets. In some embodiments, the preset targets can include aperigee altitude of 50 nautical miles, an apogee altitude of 100nautical miles, and an orbital inclination of 51.6 degrees. Theforegoing orbital parameters are provided by way of example, and inother embodiments the trajectory optimization code can be tasked withachieving other final conditions, orbits, etc.

Returning now to FIG. 19, in an orbital phase 1904 the vehicle 100 canengage in various orbital operations including, for example, transfer ofcrew and/or cargo to on-orbit stations. Typical missions on orbit couldbe from about 3 to 5 days. Once orbital operations are complete, thevehicle 100 can move to a tail-first orientation using the RCS thrusters136 described above with reference to FIGS. 1D and 1E. Next, the vehicle100 can ignite the OMS engines 122 a, b (FIG. 1A) to reduce orbitalspeed and deorbit, thereby entering a reentry phase 1905 in which thevehicle 100 descends through the Earth's atmosphere. In someembodiments, the vehicle 100 can have a weight of about 60,000 lbs. andbe at an altitude of about 400,000 ft. at a reentry start point 1905 a.In some embodiments, the angle of attack and the bank angle of thevehicle 100 can be optimized during reentry so that the vehicle 100 willbe at an altitude of about 50,000 ft. and travelling at a speed of about0.6 Mach at a reentry end point 1905 b that is about 3,000 nauticalmiles from the reentry start point 1905 a. Accelerations during thereentry phase 1905 are relatively low and can range from a maximumacceleration of from about 1.5 g's to about 2.5 g's. During a finalglide phase 1906 and subsequent landing phase 1907, the vehicle speedand descent rate can be at least generally similar to the speed anddescent of a conventional commercial jet aircraft. For example, in someembodiments the vehicle 100 can land at a speed of from about 120 mph toabout 160 mph, or about 140 mph, and at a sink rate of about 10 feet persecond. Additionally, the vehicle 100 can land on a standard runway1580. Although the foregoing reentry and landing parameters of someembodiments are provided herein by way of example, in other embodimentsthe vehicle 100 and missions thereof can have other reentry and landingparameters.

In the event the vehicle 100 lands at a runway that does not have asuitable launch sled, the vehicle 100 can be moved to a runway that doeshave a launch sled using a number of different methods. For example, insome embodiments the vehicle can be towed through the air to a newrunway by a tow aircraft. In other embodiments, the vehicle 100 caninclude provisions for jet engines that can be temporarily installed onthe wings 104 a, b (FIG. 1A) to enable the vehicle 100 to fly to the newrunway under its own power.

If an irrecoverable emergency arises during any phase of flight, thevehicle 100 can execute an abort phase 1908. In the abort phase 1908,the crew cabin 112 separates from the rest of the vehicle 100 asdescribed above with reference to FIG. 3A. Immediately after separation,the high-power escape thrusters 360 a, b propel the crew cabin 112safely away from the rest of the vehicle 100, and the recovery chute 364is deployed so that the crew cabin 112 can descend to a safe landing.The emergency parachute landing system is configured to bring the crewcabin 112 down safely on land or in water, and the cabin 112 hasprovisions that permit crew survival for an extended period of time ifrescue is delayed.

FIG. 21 is a chart listing example types of mission aborts and enginefailures/degradations that the vehicle 100 could experience, inaccordance with embodiments of the present technology. In someembodiments, there are four main types of mission aborts and two maintypes of engine failures. The least impactful of the aborts is a “ToOrbit Abort” 2102. This abort condition arises when the vehicle 100 willachieve orbit safely but will fall short of target performance. Tosuccessfully abort the mission under this condition, the vehicle 100will continue the mission until orbit is achieved, and then missioncontinuation and landing options will be assessed. A slightly moreimpactful abort is a “Once Around Abort” 2104. This abort conditionarises when the vehicle 100 will not achieve a stable orbit. To abortthis condition, the vehicle 100 will perform one orbit around the Earthand return safely to the launch site. The second most impactful abort isa “Down Range Abort” 2106. This abort condition arises when the vehicle100 is unable to perform one orbit around the Earth as required by theOnce Around Abort 2104. This abort sequence calls for the vehicle 100 toperform an emergency landing at one of a preset downrange emergencylanding sites (e.g., a landing strip or runway). The most impactfulabort is a “Return to Launch Site Abort” 2108. This abort conditionarises when the vehicle performance falls short of the conditions forthe Down Range Abort 2106. In a Return to Launch Site Abort, the vehicle100 will immediately return to and land at the launch site. In someembodiments, the propellant tanks on the vehicle 100 can include one ormore drain valves configured to rapidly discharge the propellant fromthe tanks prior to landing to ensure that the vehicle 100 lands withempty, or near empty propellant tanks during an abort. Landing withempty, or near empty propellant tanks enables the landing gear 126 a-cto be substantially lighter than would otherwise be required for landingthe vehicle 100 with full, or near full propellant tanks.

In some embodiments, the most impactful of the engine failures is an“Outboard Engine Failure” 2110 resulting from the performance of one orboth of the outboard engines 120 a, c (FIG. 1A) degrading to the pointof failure. To address this issue, the vehicle 100 executes a sequencethat commands both the outboard engines 120 a, c to shut down and thecenter engine 120 b to increase throttle to maximum thrust. The leastimpactful of these failures is an “Outboard Engine PerformanceDegradation” 2112. This failure condition is due to the performance ofone of the outboard engines 120 a or 120 c degrading, but not to thepoint of the outboard engine failure condition 2110. To address thisissue, the vehicle 100 executes a sequence that commands the opposingoutboard engine to throttle down to match the failed engine'sperformance. Throttling the opposing engine in this manner balances thethrust from the outboard engines 120 a, c and avoids exceeding the yawcontrol limit of the vehicle 100.

FIG. 22 is a flow diagram of a routine 2200 for executing an outboardengine anomaly abort sequence, in accordance with embodiments of thepresent technology. In block 2202, the routine starts when an anomaly inperformance of one of the outboard engines 120 a, c (FIG. 1A) isdetected. If the anomaly is an engine failure, then the routine proceedsto block 2204 and performs the outboard engine failure sequencedescribed above with reference to FIG. 21. Conversely, if an engineperformance degradation is detected in one of the outboard engines, thenthe routine proceeds to block 2206 and performs the outboard engineperformance degradation sequence described above with reference to FIG.21. In either situation, after execution of the appropriate sequence,the routine proceeds to block 2208. In block 2208, the vehicle initiatesthe To Orbit Abort sequence 2102 of FIG. 21 and, if the To Orbit Abortcondition is met (e.g., the vehicle will achieve orbit safely, but willfall short of target performance), then the routine proceeds to block2210 and the vehicle 100 executes the To Orbit Abort sequence outlinedin FIG. 21 (success). Conversely, if the To Orbit Abort condition is notmet (failure), then the routine proceeds to block 2212. In block 2212,the vehicle initiates the Once Around About sequence 2104 of FIG. 21and, if the Once Around Abort condition is met, then the vehicle 100executes the appropriate abort sequence (e.g., perform one orbit aroundEarth and return safely to the launch site) (success). If the OnceAround Abort condition is not met (failure), then the routine proceedsto block 2214. In block 2214, the vehicle initiates the Down Range Abortsequence 2106 of FIG. 21 and, if the Down Range Abort condition is met,then the vehicle 100 executes the appropriate abort sequence (success).Conversely, if the Down Range Abort condition is not met (failure), thenthe routine proceeds to block 2216 and performs the Return to LaunchSite Abort sequence 2108 of FIG. 21 and executes the appropriate abortsequence (e.g., immediately return to and land at the launch site).

The flow diagram of FIG. 22 and the other flow diagrams described hereindepict processes used in some embodiments of the present technology.These flow diagrams may not show all functions or exchanges of data, butinstead they provide an understanding of commands, data, and/orinformation exchanged under some embodiments of the systems. Those ofordinary skill in the relevant art will recognize that some functions orexchange of commands and data may be repeated, varied, omitted, orsupplemented, and other (less important) aspects not shown in the flowdiagrams may be readily implemented. Each of the steps depicted in theflow diagrams described herein can itself include a sequence ofoperations that need not be described herein. Those of ordinary skill inthe art can create source code, microcode, program logic arrays, etc. orotherwise implement the disclosed technology based on the flow diagramsand the Detailed Description provided herein. Such routines arepreferably stored in non-volatile memory, e.g., the memory 144 thatforms part of the vehicle controller 140 (FIG. 1A).

FIG. 23A is a block diagram of a suitable computing environment 2300 ain which the vehicle controller 140 can implement the various sequencesand routines described in detail above. In the illustrated embodiment,the vehicle 100 includes a display 2310 and a user interface 2312 whichare operably connected to the controller 140. The display 2310 caninclude one or more conventional display devices (e.g., LCD displays,LED displays, etc.) for providing textual, graphical, and otherinformation to users (e.g., vehicle crew). The user interface 2312 caninclude any suitable user interface devices and tools including, forexample, touchscreens, keyboards, keypads, joy sticks, graphical userinterfaces, etc. In one aspect of the present technology, the vehiclecontroller 140 can include or access a number of on-board softwareapplications. For example, in the illustrated embodiment the environment2300 a includes a guidance, navigation, and control (GNC) application2302, a systems management (SM) application 2304, and a vehicle checkout (VCO) application 2306. The GNC application 2302 is configured todetermine flight parameters, such as the position, velocity, andattitude of the vehicle 100 during flight. The GNC application 2302 alsoreceives and manages various outputs from vehicle sensors (e.g.,airspeed sensors, altitude sensors, acceleration sensors, pressuresensors, temperature sensors, etc.) and displays the output values tothe vehicle crew via the display 2310 and to ground control 1500 via oneor more associated displays. In addition, the GNC application 2302 alsomanages the majority of subsystems aboard the vehicle, such as theavionic subsystems, throughout the entirety of the vehicle flight fromafter sled separation to vehicle landing. The SM application 2304manages and controls the remainder of the vehicle subsystems that arenot controlled by the GNC application 2302, such as the payloadsubsystems, etc. Additionally, the SM application 2304 is alsoconfigured to identify anomalies/errors mid-flight and display them toboth the vehicle crew and the ground control crew. The VCO application2306 manages and controls all subsystems (e.g., the avionics subsystems)during their initialization process (which can occur during the vehiclepropellant loading sequence described above). The VCO application 2306also performs all of the ground and in-flight checkouts for the vehiclesystems and subsystems including, for example, determining if safeseparation conditions are met, if engine performance conditions are met,etc. The VCO application 2306 also processes ground control commandswhen the vehicle is connected to the propellant management system 1300(FIG. 13). The vehicle controller 140 can receive inputs from thevehicle crew via the user interface 2312, from ground control when thevehicle is attached to the propellant management system 1300, andthrough data feedback from the GNC application 2302, the SM application2304, and/or the VCO application 2306.

By way of an example implementation of the environment 2300 a, thevehicle controller 140 utilizes the GNC application 2302, the SMapplication 2304, and the VCO application 2306 to perform and executethe mission tasks for the vehicle 100 as described in detail above. Thesystem applications 2302, 2304 and 2306 process these tasks (2308) and,when the processing is successful, a task is executed (2314) and datafeedback is sent back to the vehicle controller 140. Conversely, whenthe processing results in a failure, the task is corrected (2316) andreperformed until successful. The data feedback can be displayed on boththe vehicle user interface 2312 or display 2320, as well as one or moreuser interfaces associated with ground control 1500.

FIG. 23B is a block diagram of a suitable computing environment 2300 bin which the sled controller 440 can implement one or more of the sledroutines described in detail above. In the illustrated embodiment, thesled controller 440 includes and/or accesses a number of softwareapplications, including, e.g., a sled control (SC) application 2320, asled systems management (SSM) application 2322, and a sled checkout(SCO) application 2324. In operation, the sled controller 440 receivesinputs from the vehicle crew via the vehicle controller 140. The sledcontroller 440 can also receive inputs from ground control 1500 via thevehicle controller 140 when the vehicle 100 is attached to thepropellant management system 1300 (FIG. 13). The sled controller 440 canalso receive data feedback from the SC application 2320, the SSMapplication 2322, and the SCO application 2324. The sled controller 440uses the three software applications to perform specific tasks inaccordance with input and commands received from the vehicle computer140 and/or ground control 1500. Such tasks can include, for example,igniting the sled engines 404 a-c, moving the support arms 416 a-c,controlling sled propellant flow, etc. These tasks are processed (2326)and, if the processing is successful, the tasks are executed (2330) andfeedback data is sent back to the sled controller 440. Alternatively, ifthe task processing results in a failure, the task is corrected (2328)and reperformed until it has been successfully executed. Data feedbackto the sled controller 440 can also be sent to the vehicle controller140 and displayed to both the vehicle crew (via, e.g., the display 2310)and ground control 1500 when the sled 400 is attached to the propellantmanagement system 1300.

FIG. 24A is a schematic diagram of the vehicle LOX tank 242 (FIG. 2),and FIG. 24B is a graph illustrating various pressures associated withthe LOX tank 242 as a function of time, in accordance with embodimentsof the present technology. Oxygen is gaseous during normal vehicleoperating conditions and liquid at temperatures below its Normal BoilingPoint (NPB) of −182.96 degrees C. Since oxygen boils at such a lowtemperature, it has a high vapor pressure at normal vehicle operatingconditions and is typically kept in heavy, round tanks capable ofwithstanding relatively high pressures (e.g., pressures over 20 psig) inconventional launch vehicles. In one aspect of the present technology,however, the LOX in the vehicle 100 is subcooled to a temperature of,e.g., −195.79 degrees C., or about −196 degrees C., to reduce the vaporpressure to less than 4 psig (e.g., 2-3 psig) across the walls of theLOX tank 242. As a result of this relatively low pressure, thefuselage/wing LOX tank 242 of the vehicle 100 can be shaped foraerodynamic efficiency and designed to support flight loads, withoutrequiring a rounded, pressure stabilized design as typically found inconventional launch vehicle LOX tanks. This feature also enables the useof lightweight composite materials, which lowers the structural tankweight, and provides a higher LOX load than NPB LOX because the LOXdensity increases with the lower temperature. The increased LOX loadimproves vehicle performance by increasing the amount of propellant thatcan be loaded into a given tank volume, improving the mass fraction ofthe flight vehicle 100.

Referring to FIG. 24A, in the illustrated embodiment the LOX tank 242can be vented outward through a first relief valve 2472 during vehicleascent. Additionally, the LOX tank 242 can also include a second reliefvalve 2474 that enables the LOX tank 242 to vent inward from theatmosphere to prevent negative pressure tank collapse on reentry.Desiccant canisters 2476 can be installed at the vent inlet to preventingestion of moisture during ground operations. During groundoperations, the LOX will be stored in the LOX tank 242 at the lowestvapor pressure possible (or approaching the lowest vapor pressurepossible), resulting in a pressure differential across the tank walls ofless than 3 psig. During engine operation, the LOX tank 242 can receivepressurization gas from a heat exchanger 2470 coupled to one or more ofthe main engines 120 a-c. A boost pump (not shown in FIG. 24A) can beused to increase the LOX pressure to, for example, about 40 psia to meetinlet conditions for the turbo pumps associated with the engines 120a-c.

Before takeoff, the LOX tank 242 can be filled and pressurized usingground-based sources of subcooled LOX (e.g., the propellant managementsystem 1300 of FIG. 13). By way of example, in some embodiments the LOXtank 242 can receive LOX at a temperature of about −196 degrees C.(i.e., about −320 degrees F.) from a LOX densification unit. Duringengine operation, pressure can be maintained in the LOX tank 242 byadding vaporized propellant gasses supplied by the heat exchanger 2470on the main engine 120. In some embodiments, the pressure differentialacross the tank walls will be maintained in a range from about 2 to 3.5psig during ascent to prevent boiling of the LOX, and to maintain tankpressure above local ambient (i.e., 14.7 psia or less) to prevent tankbuckling due to a negative pressure differential. During ascent and onorbit, residual ullage gasses vent through the first relief valve 2472.During reentry, the LOX tank 242 vents to the atmosphere via the secondrelief valve 2474 to prevent negative pressure tank collapse.

Operation of the LOX tank 242 as described above is reflected by thegraph 2478 shown in FIG. 24B. In the graph 2478, pressure in psi ismeasured along a vertical axis 2480, and time in seconds is measuredalong a horizontal axis 2482. A first plot line 2484 illustrates theatmospheric pressure during the ascent phase of flight, a second plotline 2486 illustrates tank internal pressure (absolute pressure) duringthis phase of flight, and a third plot line 2487 illustrates tankdifferential pressure during this phase of flight. As the first plotline 2484 illustrates, the atmospheric pressure drops from about 14.7psia at launch to essentially zero during ascent. As shown by the secondplot line 2486, the tank internal pressure follows this curve relativelyclosely to maintain a positive pressure differential of about 3.5 psigat all times during the launch and ascent phases of flight, as shown bythe third plot line 2487.

FIG. 25A is a schematic diagram of the left-wing fuel tank 240 a (FIG.2), and FIG. 25b is a graph 2596 illustrating various pressuresassociated with the fuel tank 240 a during vehicle ascent, in accordancewith embodiments of the present technology. Although the foregoingdescription refers to the left-wing fuel tank 240 a, it will beunderstood that the description applies equally to the right-wing fueltank 240 b. Referring to FIG. 25A, as noted above in some embodimentsthe vehicle 100 will use Jet-A as the fuel for the vehicle main engines120 a-c. Jet fuel is abundant, inexpensive, and has a naturally lowvapor pressure across vehicle operating conditions which allows it to bestored inside aerospace vehicle wings, fuselages, etc. of virtually anyshape. In the illustrated embodiment, the fuel tank 240 a includes afirst relief valve 2592, a second relief valve 2594, and a boost pump2590. As with the LOX tank 242 described above with reference to FIG.24A, during engine operation the fuel tank 240 a can maintain positivepressure by the addition of vaporized propellant gasses supplied from aheat exchanger on one or more of the vehicle engines 120 (not shown). Inorbit, residual ullage gasses can be vented from the fuel tank 240 athrough the first relief valve 2592. During reentry, the fuel tank 240 acan be vented to atmosphere via the second relief valve 2594 to preventnegative pressure collapse. Although not shown, desiccant canisters canbe installed at the vent inlet to the relief valve 2594 to preventingestion of moisture during ground operations.

Referring next to FIG. 25B, pressure in psi is measured along a verticalaxis 2595, and time in seconds is measured along a horizontal axis 2597.A first plot line 2591 illustrates the atmospheric pressure as afunction of time during vehicle ascent, and a second plot line 2593illustrates the internal (absolute) pressure of the fuel tank 240 aduring this phase of flight. The differential tank pressure across thetank wall is illustrated by a third plot line 2598. In this embodiment,the first relief valve 2592 is open and the fuel tank 240 a vents sothat the internal pressure (second plot line 2593) is essentiallyequivalent to the atmospheric pressure (first plot line 2591) for theinitial portion of vehicle ascent. At about 47 seconds after takeoff,the first relief valve 2592 closes. At this time, the vehicle will be atan altitude of about 13,000 ft. and the atmospheric pressure will beabout 9 psia. As vehicle ascent continues, the tank internal pressure isallowed to build relative to the atmospheric pressure as shown by acomparison of the first plot line 2591 to the second plot line 2593. Asa result of this pressure differential, the fuel tank 240 a has apositive pressure differential across the tank walls of from about 3psig to about 1.4 psig at the end of the vehicle ascent phase, as shownby the third plot line 2598.

FIGS. 26A and 26B are top and bottom isometric views, respectively ofthe vehicle 100. In some embodiments, the vehicle primary structure,external surfaces, etc. (including, for example, the wings 104 a, b, thefuselage 102, the vertical stabilizers 110 a, b, etc.) can beconstructed from lightweight, durable composite materials. The compositematerials can include graphite and an epoxy matrix including, forexample, a polyurethane that is resistant to micro cracking and oxygeninfusion. In some embodiments, portions of vehicle 100 can also beconstructed from metal including, for example, aluminum, titanium,stainless steel, etc. Additionally, portions of the vehicle 100 can becovered with a thermal protection system (TPS) to prevent structuraldegradation due to aerothermal heating during vehicle reentry. Forexample, in some embodiments the TPS can be applied to at least afuselage nose portion 2602, wing leading edge portions 2606 a and 2606b, and an underside portion 2604 of the fuselage 102 and the wings 104a, b. Various suitable materials known in the art may be used for theTPS. By way of example, in some embodiments the TPS can include aToughened Uni-piece Fibrous Reinforced Oxidation-resistant Compositematerial known as “TUFROC.” TUFROC can survive extreme heat environmentsup to 3,600 degrees F. and above. In other embodiments, other types ofTPS materials can be used to protect vehicle 100 from aerothermalheating degradation.

There are a number of advantages associated with embodiments of therocket powered launch sled 400 described above. For example, the launchsled 400 can enable the flight vehicle 100 to gain more velocity duringthe takeoff run than could otherwise be achieved by the unassistedvehicle 100. Additionally, by transferring propellant to the vehicle 100during the takeoff run, the sled 400 enables the vehicle 100 to leavethe ground full (or at least nearly full) of propellant, therebyminimizing (or at least reducing) the vehicle dry weight penalty andmaximizing (or at least increasing) available flight performance. Inthis way, the sled 400 can be viewed as a “first stage” of the vehicle100 that never leaves the ground. The sled 400 can also provide ameasure of safety during the critical first seconds of vehicle mainengine firing, because the sled braking system 455 (FIG. 4A) isconfigured to slow the vehicle 100 to a stop and/or hold it in place ifanomalous performance of the main vehicle engines 120 a-c is detected atany time during takeoff. Additionally, since the vehicle 100 takes offon the sled 400 and not its own landing gear 126 a-c, the landing gear126 a-c do not have to be sized to withstand the structural loadsassociated with takeoff with a full load of propellant. Instead, thelanding gear 126 a-c need only be sized to carry the lower loadsassociated with landing with empty, or near empty, propellant tanks.Accordingly, the sled 400 also eliminates the need for a heavy-dutytakeoff-rated landing gear system, thereby saving flight vehicle weightand reducing system complexity.

The above Detailed Description of examples and embodiments of thepresent technology is not intended to be exhaustive or to limit thedisclosed technology to the precise forms disclosed above. Whilespecific examples for the present technology are described above forillustrative purposes, various equivalent modifications are possiblewithin the scope of the disclosed technology, as those skilled in therelevant art will recognize. For example, while processes, routines,and/or blocks are presented in a given order, alternativeimplementations may perform processes and routines having steps, oremploy systems having blocks, in a different order, and some processes,routines or blocks may be deleted, moved, added, subdivided, combined,and/or modified to provide alternative or sub-combinations. Each of theprocesses or blocks may be implemented in a variety of different ways.Also, while processes or blocks are at times shown as being performed inseries, these processes or blocks may instead be performed orimplemented in parallel, or may be performed at different times.

Several embodiments of the present technology may take the form ofcontroller- or computer-executable instructions, such as routinesexecuted by the vehicle controller 140 and/or the sled controller 440,or by another data processing device, e.g., an onboard computer,special-purpose computer, server computer, personal computer, etc. Thoseskilled in the relevant art will appreciate that aspects of the presenttechnology can be practiced with other computer/controller systems,including other communications, data processing, or computer systemconfigurations. Aspects of the present technology can be embodied in aspecial purpose computer or data processor that is specificallyprogrammed, configured, or constructed to perform one or more of thefunctions, methods, and/or computer-executable instructions explainedherein. Accordingly, the terms “computer” and “controller” as generallyused herein refer to any data processor and can include onboard andremote computers, Internet appliances and hand-held devices (includingpalm-top computers, wearable computers, cellular or mobile phones,multi-processor systems, processor-based or programmable consumerelectronics, network computers, mini computers and the like).Information handled by these computers can be presented at any suitabledisplay medium, including a liquid crystal display.

EXAMPLES

The following examples provide additional embodiments of the presenttechnology:

1. A method of operating a reusable space vehicle, the methodcomprising:

-   -   releasably attaching the space vehicle to a launch sled;    -   igniting one or more rocket engines on the launch sled;    -   igniting one or more rocket engines on the space vehicle; and    -   providing propellant from one or more propellant tanks on the        launch sled to the one or more rocket engines on the space        vehicle and to the one or more rocket engines on the launch        sled, while the one or more rocket engines on the space vehicle        and the one or more rocket engines on the launch sled provide        thrust for launch of the space vehicle.

2. The method of example 1, further comprising:

-   -   accelerating the launch sled toward a takeoff speed of the space        vehicle;    -   at or near the takeoff speed, ceasing to provide propellant from        the one or more propellant tanks on the launch sled to the one        or more rocket engines on the space vehicle; and    -   releasing the space vehicle from the launch sled.

3. The method of example 2, further comprising, at or near the takeoffspeed, providing propellant from one or more propellant tanks on thespace vehicle to the one or more rocket engines on the space vehicle.

4. The method of example 2, wherein the space vehicle is configured tofly into space, orbit the Earth, and then reenter the Earth'satmosphere, and wherein the method further comprises:

-   -   flying the vehicle toward a landing site;    -   deploying a landing gear on the vehicle; and    -   horizontally landing the vehicle on the landing gear at the        landing site.

5. The method of example 1, further comprising returning ventedpropellant from the space vehicle to the one or more propellant tanks onthe launch sled as the one or more rocket engines on the space vehicleprovide thrust for launch of the space vehicle.

6. The method of example 1 wherein the one or more propellant tanks onthe launch sled includes an oxidizer tank, and wherein providingpropellant from the one or more propellant tanks on the launch sledincludes providing oxidizer from the oxidizer tank to the one or morerocket engines on the launch sled and to the one or more rocket engineson the space vehicle.

7. The method of example 1 wherein the one or more propellant tanks onthe launch sled includes a fuel tank, and wherein providing propellantfrom the one or more propellant tanks on the launch sled includesproviding fuel from the fuel tank to the one or more rocket engines onthe launch sled and to the one or more rocket engines on the spacevehicle.

8. The method of example 1 wherein the one or more propellant tanks onthe launch sled include an oxidizer tank and a fuel tank, and whereinproviding propellant from the one or more propellant tanks includesproviding oxidizer from the oxidizer tank and fuel from the fuel tank tothe one or more rocket engines on the space vehicle and to the one ormore rocket engines on the launch sled.

9. The method of example 1 wherein the launch sled includes a firstoxidizer tank and a second oxidizer tank, and wherein providingpropellant from the one or more propellant tanks includes providingoxidizer from the first oxidizer tank to the one or more rocket engineson the launch sled and providing oxidizer from the second oxidizer tankto the one or more rocket engines on the space vehicle.

10. The method of example 1 wherein the launch sled includes a firstfuel tank and a second fuel tank, and wherein providing propellant fromthe one or more propellant tanks includes providing fuel from the firstfuel tank to the one or more rocket engines on the launch sled andproviding fuel from the second fuel tank to the one or more rocketengines on the space vehicle.

11. The method of example 1, further comprising:

-   -   prior to igniting the one or more rocket engines on the launch        sled and the one or more rocket engines on the space vehicle,        releasably attaching a propellant umbilical between the launch        sled and the space vehicle, wherein providing propellant from        one or more propellant tanks on the launch sled to the one or        more rocket engines on the space vehicle includes flowing the        propellant through the propellant umbilical;    -   after igniting the one or more rocket engines on the launch sled        and the one or more rocket engines on the space vehicle,        accelerating the launch sled toward a takeoff speed of the space        vehicle;    -   at or near the takeoff speed, disconnecting the propellant        umbilical from the space vehicle; and    -   releasing the space vehicle from the launch sled.

12. The method of example 1 wherein releasably attaching the spacevehicle to the launch sled includes releasably coupling one or moresupport arms extending from the launch sled to the vehicle, and whereinthe method further comprises:

-   -   supporting the space vehicle solely by the one or more support        arms.

13. The method of example 12, further comprising rotating the one ormore support arms relative to the launch sled to increase an angle ofattack of the space vehicle.

14. The method of example 1, further comprising:

-   -   prior to releasably attaching the space vehicle to the launch        sled, rolling the vehicle onto the launch sled on a vehicle        landing gear, wherein releasably attaching the space vehicle to        the launch sled includes releasably coupling one or more support        arms extending from the launch sled to the vehicle; and    -   retracting the vehicle landing gear so that the space vehicle is        supported solely by the support arms.

15. The method of example 14, further comprising, prior to retractingthe vehicle landing gear, rotating the one or more support arms relativeto the launch sled to raise the vehicle off of the landing gear.

16. A space vehicle system, comprising:

-   -   a reusable space vehicle having one or more rocket engines; and    -   a launch sled having one or more rocket engines, wherein the        launch sled is configured to support the space vehicle during        launch from Earth, and wherein the launch sled is further        configured to provide propellant to the one or more rocket        engines of the launch sled and to the one or rocket engines of        the space vehicle during launch of the space vehicle.

17. The space vehicle system of example 16 wherein the reusable spacevehicle is a horizontal takeoff/horizontal landing (HTHL) spaceplane.

18. The space vehicle system of example 16 wherein the reusable spacevehicle is a single-stage-to-orbit (SSTO) spaceplane.

19. The space vehicle system of example 16 wherein the one or morerocket engines on the vehicle are bipropellant engines that use liquidoxygen and liquid fuel.

20. The space vehicle system of example 16:

-   -   wherein the launch sled includes a first propellant tank,    -   wherein the space vehicle includes a second propellant tank that        provides propellant to the one or more rocket engines on the        space vehicle, and    -   wherein the launch sled further includes a propellant umbilical        configured to transfer propellant from the first propellant tank        to the second propellant tank during launch of the space        vehicle.

21. The space vehicle system of example 20 wherein the first propellanttank provides propellant to the one or more rocket engines on the launchsled.

22. The space vehicle system of example 16:

-   -   wherein the launch sled includes a first propellant tank and a        second propellant tank,    -   wherein the first propellant tank provides propellant to the one        or more rocket engines on the launch sled,    -   wherein the space vehicle includes a third propellant tank that        provides propellant to the one or more rocket engines on the        space vehicle, and    -   wherein the launch sled further includes a propellant umbilical        configured to transfer propellant from the second propellant        tank to the third propellant tank during launch of the space        vehicle.

23. The space vehicle system of example 16 wherein the launch sledincludes a first oxidizer tank, wherein the space vehicle includes asecond oxidizer tank that provides oxidizer to the one or more rocketengines on the space vehicle, and wherein the launch sled is configuredto provide oxidizer from the first oxidizer tank to the second oxidizertank during launch of the space vehicle.

24. The space vehicle system of example 16 wherein the launch sledincludes a first fuel tank, wherein the space vehicle includes a secondfuel tank that provides fuel to the one or more rocket engines on thespace vehicle, and wherein the launch sled is configured to provide fuelfrom the first fuel tank to the second fuel tank during launch of thespace vehicle.

25. The space vehicle system of example 16:

-   -   wherein the launch sled includes one or more support arms        configured to releasably support the space vehicle on the launch        sled,    -   wherein a distal end portion of each of the one or more support        arms includes a hold and release mechanism configured to        releasably attach to a corresponding interface on the space        vehicle, and    -   wherein each of the one or more support arms is movable to        change the position of the space vehicle relative to the launch        sled.

26. The space vehicle system of example 16:

-   -   wherein the launch sled includes a first propellant tank,    -   wherein the space vehicle includes a second propellant tank that        provides propellant to the one or more rocket engines on the        space vehicle,    -   wherein the launch sled further includes a propellant umbilical        configured to transfer propellant from the first propellant tank        to the second propellant tank during launch of the space        vehicle, the propellant umbilical extending between a propellant        outlet interface on an upper portion of the launch sled to a        propellant inlet interface on an underside of the space vehicle,    -   wherein the launch sled further includes at least one support        arm having a proximal end portion pivotally coupled to the upper        portion of the launch sled laterally adjacent to the propellant        outlet interface, and a distal end portion configured to be        releasably coupled to an interface fitting on the underside of        the space vehicle laterally adjacent to the propellant inlet        interface, and    -   wherein the at least one support arm is rotatable to change the        position of the space vehicle relative to the launch sled.

27. A reusable space vehicle, comprising:

-   -   a pair of wings configured to provide lift during flight of the        space vehicle in the Earth's atmosphere;    -   one or more rocket engines; and    -   an oxidizer tank configured to provide liquid oxygen to the one        or more rocket engines, the oxidizer tank having a        non-cylindrical and non-spherical shape.

28. The reusable space vehicle of example 27, further comprising afuselage having an external sidewall, and wherein the external sidewallforms a portion of the oxidizer tank.

29. The reusable space vehicle of example 27 wherein the oxidizer tankis configured to withstand an internal pressure of less than 4 psig.

30. The reusable space vehicle of example 27 wherein the oxidizer tankis configured to carry liquid oxygen at a temperature of from about −196degrees C. to about −182 degrees C. during launch of the space vehicle.

31. The reusable space vehicle of example 27, further comprising one ormore structural interfaces on an underside thereof configured toreleasably attach the space vehicle to a rocket-powered launch sled fortakeoff of the space vehicle.

32. The reusable space vehicle of example 27, further comprising:

-   -   one or more structural interfaces on an underside thereof        configured to releasably attach the space vehicle to a        rocket-powered launch sled for takeoff of the space vehicle; and    -   a landing gear that is only used for landing the space vehicle.

33. A launch sled for launching a reusable space vehicle, the spacevehicle having one or more rocket engines, the launch sled comprising:

-   -   one or more rocket engines for providing thrust to the launch        sled during launch of the vehicle; and    -   a propellant tank configured to be operably coupled in fluid        communication with the space vehicle, wherein the propellant        tank is configured to provide propellant to the one or more        engines of the space vehicle during launch of the space vehicle.

34. The launch sled of example 33, further comprising a propellantumbilical configured to be releasably connected between the launch sledand the space vehicle to transfer propellant therebetween during launchof the space vehicle.

35. The launch sled of example 33, further comprising a plurality ofmovable support arms configured to releasably attach the space vehicleto the launch sled and move the space vehicle relative to the launchsled.

36. A method for loading liquid oxygen into a horizontaltakeoff/horizontal landing space vehicle, the method comprising:

-   -   cooling the liquid oxygen to a temperature of about −196        degrees C. or less; and    -   flowing the cooled liquid oxygen into an oxidizer tank on the        space vehicle.

37. The method of example 36 wherein cooling the liquid oxygen includesflowing the liquid oxygen through a liquid nitrogen heat exchanger, andwherein the method further comprises:

-   -   venting oxygen from the oxidizer tank;    -   recirculating the vented oxygen through the liquid nitrogen heat        exchanger to re-cool the liquid oxygen to a temperature of about        −196 degrees C. or less; and    -   flowing the re-cooled liquid oxygen back to the oxidizer tank.

38. The method of example 36 wherein the space vehicle is mounted to alaunch sled having a first oxidizer tank, wherein the oxidizer tank onthe space vehicle is a second oxidizer tank, and wherein the methodfurther comprises flowing the cooled liquid oxygen into the firstoxidizer tank on the sled and then flowing the cooled liquid oxygen fromthe first oxidizer tank into the second oxidizer tank.

39. The method of example 36 wherein flowing the liquid oxygen into theoxidizer tank includes flowing the liquid oxygen into an oxidizer tankhaving a non-cylindrical, non-spherical shape.

40. The method of example 36 wherein flowing the liquid oxygen into theoxidizer tank includes flowing the liquid oxygen into an oxidizer tankon a space vehicle having a pair of wings for horizontal takeoff andlanding in the Earth's atmosphere.

41. The method of example 36 wherein flowing the liquid oxygen into theoxidizer tank includes flowing the liquid oxygen into an oxidizer tankin a fuselage of a space vehicle having a pair of wings for horizontaltakeoff and landing in the Earth's atmosphere, the oxidizer tank havinga non-cylindrical, non-spherical shape.

References throughout the foregoing description to features, advantages,or similar language do not imply that all of the features and advantagesthat may be realized with the present technology should be or are in anysingle embodiment. Rather, language referring to the features andadvantages is understood to mean that a specific feature, advantage, orcharacteristic described in connection with an embodiment is included inat least one embodiment of the present technology. Thus, discussion ofthe features and advantages, and similar language, throughout thisspecification may, but do not necessarily, refer to the same embodiment.Furthermore, the described features, advantages, and characteristics ofthe present technology may be combined in any suitable manner in one ormore embodiments. One skilled in the relevant art will recognize thatthe present technology can be practiced without one or more of thespecific features or advantages of a particular embodiment. In otherinstances, additional features and advantages may be recognized incertain embodiments that may not be present in all embodiments of thepresent technology.

Any patents and applications and other references noted herein,including any that may be listed in accompanying filing papers, areincorporated herein by reference. To the extent that any materialsincorporated herein by reference conflict with the present disclosure,the present disclosure controls. Aspects of the present technology canbe modified, if necessary, to employ the systems, functions, andconcepts of the various references described above to provide yetfurther implementations of the present technology.

While the above description describes various embodiments of thedisclosed technology and the best mode contemplated, regardless howdetailed the above text, the technology can be practiced in many ways.Details of the system may vary considerably in its specificimplementation, while still being encompassed by the present disclosure.As noted above, particular terminology used when describing certainfeatures or aspects of the disclosed technology should not be taken toimply that the terminology is being redefined herein to be restricted toany specific characteristics, features, or aspects of the disclosedtechnology with which that terminology is associated. In general, theterms used should not be construed to limit the disclosed technology tothe specific examples disclosed in the specification, unless the aboveDetailed Description section explicitly defines such terms. As usedherein, the term “and/or”, as in “A and/or B,” refers to A alone, Balone, and both A and B. From the foregoing, it will be appreciated thatspecific embodiments of the disclosed technology have been describedherein for purposes of illustration, but that various modifications maybe made without deviating from the spirit and scope of the presentdisclosure.

From the foregoing, it will be appreciated that specific embodiments ofthe invention have been described herein for purposes of illustration,but that various modifications may be made without deviating from thespirit and scope of the various embodiments of the invention. Further,while various advantages associated with certain embodiments of theinvention have been described above in the context of those embodiments,other embodiments may also exhibit such advantages, and not allembodiments need necessarily exhibit such advantages to fall within thescope of the invention. Accordingly, the invention is not limited,except as by the appended claims.

Although certain aspects of the invention are presented below in certainclaim forms, the applicant contemplates the various aspects of theinvention in any number of claim forms. Accordingly, the applicantreserves the right to pursue additional claims after filing thisapplication to pursue such additional claim forms, in either thisapplication or in a continuing application.

I/We claim:
 1. A method of operating a reusable space vehicle, themethod comprising: releasably attaching the space vehicle to a launchsled; igniting one or more rocket engines on the launch sled; ignitingone or more rocket engines on the space vehicle; providing propellantfrom one or more propellant tanks on the launch sled to the one or morerocket engines on the space vehicle and to the one or more rocketengines on the launch sled, while the one or more rocket engines on thespace vehicle and the one or more rocket engines on the launch sledprovide thrust for launch of the space vehicle; and returning ventedpropellant from the space vehicle to the one or more propellant tanks onthe launch sled as the one or more rocket engines on the space vehicleprovide thrust for launch of the space vehicle.
 2. The method of claim1, further comprising: accelerating the launch sled toward a takeoffspeed of the space vehicle; at or near the takeoff speed, ceasing toprovide propellant from the one or more propellant tanks on the launchsled to the one or more rocket engines on the space vehicle; andreleasing the space vehicle from the launch sled.
 3. The method of claim2, further comprising, at or near the takeoff speed, providingpropellant from one or more propellant tanks on the space vehicle to theone or more rocket engines on the space vehicle.
 4. The method of claim2, wherein the space vehicle is configured to fly into space, orbit theEarth, and then reenter the Earth's atmosphere, and wherein the methodfurther comprises: flying the vehicle toward a landing site; deploying alanding gear on the vehicle; and horizontally landing the vehicle on thelanding gear at the landing site.
 5. The method of claim 1 wherein theone or more propellant tanks on the launch sled includes an oxidizertank, and wherein providing propellant from the one or more propellanttanks on the launch sled includes providing oxidizer from the oxidizertank to the one or more rocket engines on the launch sled and to the oneor more rocket engines on the space vehicle.
 6. The method of claim 1wherein the one or more propellant tanks on the launch sled includes afuel tank, and wherein providing propellant from the one or morepropellant tanks on the launch sled includes providing fuel from thefuel tank to the one or more rocket engines on the launch sled and tothe one or more rocket engines on the space vehicle.
 7. The method ofclaim 1 wherein the one or more propellant tanks on the launch sledinclude an oxidizer tank and a fuel tank, and wherein providingpropellant from the one or more propellant tanks includes providingoxidizer from the oxidizer tank and fuel from the fuel tank to the oneor more rocket engines on the space vehicle and to the one or morerocket engines on the launch sled.
 8. The method of claim 1 wherein thelaunch sled includes a first oxidizer tank and a second oxidizer tank,and wherein providing propellant from the one or more propellant tanksincludes providing oxidizer from the first oxidizer tank to the one ormore rocket engines on the launch sled and providing oxidizer from thesecond oxidizer tank to the one or more rocket engines on the spacevehicle.
 9. The method of claim 1 wherein the launch sled includes afirst fuel tank and a second fuel tank, and wherein providing propellantfrom the one or more propellant tanks includes providing fuel from thefirst fuel tank to the one or more rocket engines on the launch sled andproviding fuel from the second fuel tank to the one or more rocketengines on the space vehicle.
 10. The method of claim 1, furthercomprising: prior to igniting the one or more rocket engines on thelaunch sled and the one or more rocket engines on the space vehicle,releasably attaching a propellant umbilical between the launch sled andthe space vehicle, wherein providing propellant from one or morepropellant tanks on the launch sled to the one or more rocket engines onthe space vehicle includes flowing the propellant through the propellantumbilical; after igniting the one or more rocket engines on the launchsled and the one or more rocket engines on the space vehicle,accelerating the launch sled toward a takeoff speed of the spacevehicle; at or near the takeoff speed, disconnecting the propellantumbilical from the space vehicle; and releasing the space vehicle fromthe launch sled.
 11. The method of claim 1 wherein releasably attachingthe space vehicle to the launch sled includes releasably coupling one ormore support arms extending from the launch sled to the vehicle, andwherein the method further comprises: supporting the space vehiclesolely by the one or more support arms.
 12. The method of claim 11,further comprising rotating the one or more support arms relative to thelaunch sled to increase an angle of attack of the space vehicle.
 13. Themethod of claim 1, further comprising: prior to releasably attaching thespace vehicle to the launch sled, rolling the vehicle onto the launchsled on a vehicle landing gear, wherein releasably attaching the spacevehicle to the launch sled includes releasably coupling one or moresupport arms extending from the launch sled to the vehicle; andretracting the vehicle landing gear so that the space vehicle issupported solely by the support arms.
 14. The method of claim 13,further comprising, prior to retracting the vehicle landing gear,rotating the one or more support arms relative to the launch sled toraise the vehicle off of the landing gear.
 15. A method for loadingliquid oxygen into a horizontal takeoff/horizontal landing spacevehicle, wherein the space vehicle is mounted to a launch sled having afirst oxidizer tank, wherein the space vehicle has a second oxidizertank, and wherein the method comprises: cooling the liquid oxygen to atemperature of about −196 degrees C. or less; flowing the cooled liquidoxygen into the first oxidizer tank on the sled; and flowing the cooledliquid oxygen from the first oxidizer tank into the second oxidizer tankon the space vehicle.
 16. The method of claim 15 wherein cooling theliquid oxygen includes flowing the liquid oxygen through a liquidnitrogen heat exchanger, and wherein the method further comprises:venting oxygen from the second oxidizer tank; recirculating the ventedoxygen through the liquid nitrogen heat exchanger to re-cool the liquidoxygen to a temperature of about −196 degrees C. or less; and flowingthe re-cooled liquid oxygen back to the first oxidizer tank.
 17. Themethod of claim 15 wherein flowing the liquid oxygen into the secondoxidizer tank includes flowing the liquid oxygen into an oxidizer tankhaving a non-cylindrical, non-spherical shape.
 18. The method of claim15 wherein flowing the liquid oxygen into the second oxidizer tankincludes flowing the liquid oxygen into an oxidizer tank on a spacevehicle having a pair of wings for horizontal takeoff and landing in theEarth's atmosphere.
 19. The method of claim 15 wherein flowing theliquid oxygen into the second oxidizer tank includes flowing the liquidoxygen into an oxidizer tank in a fuselage of a space vehicle having apair of wings for horizontal takeoff and landing in the Earth'satmosphere, the oxidizer tank having a non-cylindrical, non-sphericalshape.
 20. A method for loading liquid oxygen into a horizontaltakeoff/horizontal landing space vehicle, the method comprising: coolingthe liquid oxygen to a temperature of about −196 degrees C. or less; andflowing the cooled liquid oxygen into an oxidizer tank on the spacevehicle, wherein flowing the liquid oxygen into the oxidizer tankincludes flowing the liquid oxygen into an oxidizer tank having anon-cylindrical, non-spherical shape.
 21. The method of claim 20 whereincooling the liquid oxygen includes flowing the liquid oxygen through aliquid nitrogen heat exchanger, and wherein the method furthercomprises: venting oxygen from the oxidizer tank; recirculating thevented oxygen through the liquid nitrogen heat exchanger to re-cool theliquid oxygen to a temperature of about −196 degrees C. or less; andflowing the re-cooled liquid oxygen back to the oxidizer tank.
 22. Themethod of claim 20 wherein the space vehicle is mounted to a launch sledhaving a first oxidizer tank, wherein the oxidizer tank on the spacevehicle is a second oxidizer tank, and wherein the method furthercomprises flowing the cooled liquid oxygen into the first oxidizer tankon the sled and then flowing the cooled liquid oxygen from the firstoxidizer tank into the second oxidizer tank.
 23. The method of claim 20wherein flowing the liquid oxygen into the oxidizer tank includesflowing the liquid oxygen into an oxidizer tank on a space vehiclehaving a pair of wings for horizontal takeoff and landing in the Earth'satmosphere.
 24. The method of claim 20 wherein flowing the liquid oxygeninto the oxidizer tank includes flowing the liquid oxygen into anoxidizer tank in a fuselage of a space vehicle having a pair of wingsfor horizontal takeoff and landing in the Earth's atmosphere.